Zero-G Devices and Weightlessness Simulators (1961)

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1. DEVICES FOR PRODUCING THE ZERO-G CONDITION A. Vertical Motion Devices The simplest method to produce the zero-G condition for biomedical and physical experimentation consists in the utiliza- tion of the vertical free-fall state. If a body moves vertically downward, an upward acting force of inertia which compensates the body's weight becomes effective. The basic equations for presenting the physical conditions for a free-fall body in a vacuum are: h(height)= g t2; (2) ~2~ and, v (velocity) = g. t. (3) However, strong frictional forces from the air will soon be produced so that the free fall of the body is slowed down and its original weight restored because of these forces. Speaking in terms of kinematics, the acceleration of gravity acting on the body during its free fall cannot be maintained for an appreciable length of time, and therefore the velocity of the falling body soon approaches a constant value. In order to find the general formula which can be applied for determining the parameters pertaining to a body falling within the atmosphere, it is assumed that the total weight of the body is designated by G. The final velocity, where the weight of the body and the drag brought about by the resistance of the air are balanced, may be called the terminal velocity or the equilibrium velocity, and is designated by ve. In order to arrive at ve, we assume that G = D (4) D = (CD. p . A. ve2) (5)

(6) D. o • where CD is the drag coefficient, p represents the density of the atmosphere, and A is the area of the body in the direction of fall. Assuming drag as expressed in (5), and since £ - g - ad, (7) dt a^ represents the acceleration of the drag forces. Then = .§_•_._, (9) g \dt/ \ve/ 2 2 ve i ve dt=i:—o 2~~ - dv. (10) a V - V B e In order to find the height (h) at which G = D, we must find the integral of v. dt for the values from v to 0. Thus, v v /2 /• v'dt = e / X -dv. (H) 0 0 In solving this equation for h, we obtain h- ve2. 1n ve2 re or, 2g v * - v* (12) (13) With increasing time of fall v reaches the value v . By assuming

v to be smaller than unity, and by designating this quotient as k, ve we obtain y 1 . ve k2 = 1 - e e2 = 1 - k2 . (16) -2gh/ve2 1n (1 - k2) , (17) h= - ^_ 1n (1 - k2) . (18) 2g Equation (18) yields the sought relationship of height, gravity, and final velocity for any finite value of k; i. e. , for any condition of free fall in which the falling body has not yet reached its termi- nal velocity. Practically, a few seconds of virtual weightlessness have been obtained in free-fall experiments from drop towers. In order to produce longer free-fall periods, itwould be necessary to ascend to higher altitudes and regions of reduced air densities. Moreover, it was suggested to drop the body from a balloon, because its re- lease from an aircraft would give rise to considerable frictional forces from the air associated with relatively high decelerations. There are a lot of other difficulties involved, such as slowing down the body before impact, the recovery procedures, safety measures in case of human and animal experiments, and the relatively small yield compared with the considerable preparations required for a single experiment. If a body is subjected to a downward acceleration smaller than that of gravity, the forces of inertia associated with accelera- tion are subtracted from the gravitational force. This condition has been designated as the subgravity state. Similarly, if a force of inertia is added to the gravitational force, as for instance, dur- ing centrifugation or liftoff of a rocket, the forces involved add vectorially and produce a state of "super-gravity" or increased weight. This would also be the case during a powered approach of the earth with an acceleration larger than that of the earth's gravity.

The Drop Tower 1. The SAM dropping experiments consisted of small pay- loads dropped from the School of Aerospace Medicine Building in 1948-49. 2. A schematic picture of a drop tower is shown in Figure 1. Actually, experimental payloads can be dropped from any structure, if a free-fall condition is secured. In most cases, re- covery of the payload can be obtained by an elastic damping device. 3. The height of drop towers is limited by the resistance of the air which compensates the acceleration of gravity completely after about six seconds. Thus, only a few seconds of zero-G and subgravity can be produced in this manner. A body falling from an altitude of about 400 feet obtains its final speed (ve) of about 180 feet per second at the end of this fall. 4. Basic equations are given above. 5. During early experiments made by scientists from the School of Aviation Medicine, Randolph Air Force Base, Texas, the behavior of a burning candle during the state of weightlessness and of solidified carbon dioxide (dry ice) was filmed. The pay- loads were manually dropped from a height of about 50 feet and manually recovered by means of an elastic canvas. 6. No other instrumentation was used than that described in Paragraph 5. 7. No special safety measures necessary except protecting the recovery crew from the falling body. In case of human and animal experiments dropped from high altitudes (aircraft or bal- loons), special safety requirements are needed. 8. No actual cost figures are available. Generally, the expense of a simple drop tower is relatively low. 9. No information available. 10. Copies of the film mentioned in paragraph 5 above are available from the Department of Research Photography, School of Aerospace Medicine, Brooks Air Force Base, San Antonio, Texas.

Figure 1. Drop Tower WADC Vertical Deceleration Tower (Drop Tower) 1. The Vertical Deceleration Tower was undergoing proof testing at the Aerospace Medical Laboratory, Wright Air Develop- ment Center, Air Research and Development Command, Dayton, Ohio. The device is operated by personnel of the Escape Section, Biophysics Branch. Size of the present staff unknown. 2. Two very heavy I-beams 70 feet in height constitute the major structure components of the tower. They are mounted in a pit 20 feet deep and extend 50 feet above floor level. On each beam is mounted a precision rail which guides the test cart during its drop. This cart, which weighs 1900 pounds, is so designed that the test items such as seats, capsules, etc. , can be mounted in any orientation to the direction of the G-forces, which is parallel to the rails. This cart is dropped from a given height and falls 8

freely (for all practical purposes) to floor level where a tapered plunger, approximately 4 feet in length and 8 inches in diameter at its widest part, mounted on the bottom of the cart, enters a tube filled with water. About 2 seconds of virtual weightlessness occur during the drop. Figures 2 and 3 show the structure and location of the test subject. 3. Although the tower was designed to investigate decelera- tion effects on men and equipment only, it can also be utilized for the study of very short periods of subgravity. If the device is used in its present form, the subgravity state is uniform and stopped by a deceleration not exceeding 4 G-units of force. How- ever, by the installation of an acceleration control various degrees of subgravity should be obtained. The tower could then be used as a device to investigate the effects of different amounts of sub- gravity. 4. The acceleration-deceleration pattern depends on drop height, total cart weight, friction between rail and guide, and finally, the degree of plunger tapering. By either controlling the friction or the acceleration of the suspension cables, any sub- gravity pattern (0<a<l) can be produced. 5. Electronic and photographic instrumentation will be in- tegrated with a common time base. 6. No information available. 7. For maximum safety, only the top 50 feet of the tower is normally used for testing, the remaining 20 feet which are all be- low floor level, is reserved for safety overruns. In the event that the primary deceleration device (tapered plunger) should fail, the cart would continue to fall into the pit, where it would first engage mechanical brake shoes, and then actuate a metal cutting type of decelerator. Both of these devices are completely independent of each other, and either is capable of stopping the cart with full payload, falling from the top of the tower, with a deceleration of no more than 4 g. However, as they would normally both operate simultaneously, the emergency stopping deceleration would ap- proach 8 g, still well below the normal tolerance level of the human body. 8. As the only materials needed for operation are gravity, water, instrumentation recording material, and operating per- sonnel, the cost of testing is small. No actual figures available.

(Official USAF Photo) Figure 2. WADC Vertical Deceleration Tower 10

(Official USAF Photo) Figure 3. WADC Vertical Deceleration Tower: Location of the Subject 11

9. In view of the planned deceleration program of the Aerospace Medical Laboratory, it is expected that the facility will be utilized almost exclusively for this purpose during the next 18 months. Experiments which do not involve elaborate instru- mentation can be arranged, subject to Aerospace Medical Labora- tory's priority commitments, by request through Headquarters, WADC. At present, the Chief of the Biophysics Branch, and Mr. Charles Demsey, Chief, Escape Section, are responsible for control and direction of this facility. 10. Reference: WADC Vertical Deceleration Tower (Drop Tower). Report prepared for the NRC - Armed Forces Committee on Bioastronautics. The Elevator 1. Various types of elevators can be used for subgravity research. No information available about actual utilization. 2. Description of the mechanical systems can be obtained from companies which construct such devices. A schematic repre- sentation of the various states of subgravity obtainable in an ele- vator is given in Figure 4a and 4b. 3. The possibilities of utilizing vertical motion for our pur- pose are not bound to the free-fall condition. In elevator type locomotion, various amounts of subgravity can be produced during ascent and descent. From the laws of kinematics it follows that the acceleration must be kept constant in experiments of this kind; i. e. , one is free to begin the motion of the elevator with any initial velocity desired. In order to obtain maximum periods of subgravity, the experiment must be started with the highest velocity the elevator can attain in an upward motion. A state of subgravity of a certain amount can be produced by superimposing an upward acceleration upon the normal acceleration of gravity. Under these conditions the motion of the elevator car is such that the upward velocity will gradually be consumed by gravity, so that after having reached the maximum elevation the car starts to drop and reverses the pattern of motion in its downward course. 4. The total acceleration dv of a body in a straight upward or downward motion is given by dt if the direction of acceleration is upward. It is convenient to ex- press the acceleration a in terms of g by writing a=n.g, (20) 12

DIRECTION OF MOTION F, Fi DIRECTION OF MOTION DIRECTION OF MOTION ELEVATOR AT REST OR DURING NON-ACCELER- ATED MOTION: SINCE THE MAN IS FULLY SUPPORTED BY THE BOTTOM OF THE CAR, THE FORCE OF SRAVTTYIFg) 13 COMP- ENSATED BY EXTERNAL FORCES(F»I. HE THUS ACCELERATION DOWN- WARD AT 16 6 FT/ SEC 2 ONE HALF OF THE FORCE OF GRAVITY IFgllS COMP- ENSATED BY INERTIAL FORCES (Fi).THE OTHER HALF. BY EXTERNAL FOR- CES ( Ft) THE MAN IN THE CAR EXPERIENCES A DECELERATION UPWARD AT IC.I FT/SEC. I ONE HALF OF THE FORCE OF GRAVITY (Fg)IS COMP- ENSATED BY INERTIAL FORCES (Fi) THE OTHER HALF BY EXTERNAL FOR- CES (Ft) THE MAN IN THE CAR EXPERIENCES A CON- ACCELERATION DOWN- WARD AT 32 2 FT/SEC. • (FREE FALL). THE FORCE OF GRAVITY (Fg)IS EN- TIRELY COMPENSATED BY THE FORCE OF INERTIA (Fi) THE MAN IN THE CAR IS UNSUPPORTED AND ENT1RLEY WEIGHTLESS DECELERATION UPWARD AT 322 FT/SEC * THE THE INERTIAL FORCES (Fi)COMPENSATE ENTIRE- LY THE FORCE OF GRAV- ITY (Fg) THE MAN IN THE CAR IS WEIGHTLESS THIS SITUATION CAN BE OB- TAINED FOR VERY SHORT WEIGHT. NO INERT1AL FORCES ARE ACTIVE CONSIDERABLE REDUCT- ION COX) OF HIS WEIGHT. SIDERABLE REDUCTION (50%) OF HIS WEIGHT PERIODS ONLY Figure 4a. Sub-gravity and Zero-G in Elevators 0.4 0.6 n Grovify (From F. Haber, 1952. Epitome of Space Medicine. USAF Sch. of Aviat. Med.) Figure 4b. The Duration of Subgravity as Obtainable in Elevators. 13

thus transforming Equation (19) into S|Y. = n.g - g = - (1 - n).g. (21) If n is equal to zero the gravity free state results, and if o<n<l the subgravity state prevails. The solution of Equation (20) is v= vo - (1 - n).g.t. (22) For t = 0, v equals the initial speed VQ. With increasing time, v reaches the value zero: v= o= vo - (1 - n).g. to , (23) to = . vo . (24) (1 - n)'g After this time the speed becomes directed downward and reaches the value -vo. The pertinent time is (25) (1 - n)'g The elevation of the body as a function of time is the integral of Equation (22) and is h= vo. t - 1 - n . g. t2. (26) 2 The height is zero for t - 0 and becomes zero again for h = o = Vn'*d - 1 - n . g. t2d . (27) 2 td_ 2vo , (28) (1 - n)g which is identical to Equation (25). This equation gives the dura- tion t(j as a function of gravity and initial speed. For some purposes it is convenient to have the duration as a function of gravity and elevation. This is achieved by introducing Equation (23) in Equa- tion (26) and obtaining h = vo2 . (29) 2(1 - n).g 14

By this equation the speed vo in Equation (25) can be replaced by the elevation h yielding to td «/ 8 .h . (30) F (1 - n). g Equations (25) and (30), giving the relation of elevation, gravity, initial speed, and height, are plotted in Figure 4. 5. No data available. Elevator car can be instrumented to measure physiological variables and performance characteristics of test subjects, behavior of materials and substances, of animals, etc. 6. See Paragraph 5. 7. Safety factors and requirements in accordance with regu- lations for operating elevators. 8. Since elevators of various types are available (the use of elevators in the Empire State Building, the Carlsbad Caverns, and mine shafts, were previously considered), costs may have to cover the operating costs of these devices. Special safety equipment and scientific instrumentation optional. 9. No information available. 10. References included in the general bibliography attached to this report. The Subgravity Tower 1. This is a device for producing short periods of increased and decreased weight to study their physiological and psychological effects. The device is available at the Aviation Medicine Research Center in Rome, Italy. A schematic picture is given in Figure 5. 2. The Subgravity Tower consists of a metal-trestle tower, about 14 m high. Inside of this structure is a platform weighing about 50 kg which, in turn, holds the subject, medical instru- mentation, film camera, and ballast to compensate for weight differences among subjects. The following parts are necessary for proper functioning of the device: 15

a. Power device: Consists of four bundles of elastic (rubber) cords. When they are brought under tension by a mechanical winding procedure, a certain amount of potential energy is available for propelling the plat- form upward. b. Release device: Consists of a set of levers which arrest the platform at the base of the structure. The rubber cables lift the platform upon release. c. Braking device: Consists of two vertical steel cables which diverge at the upper end and exert an effective braking action without any jerks. d. Recovery device: Allows arrest of the compartment or platform upon completion of motion and to engage it to the launching device. e. G-control: Consists of an accelerometer attached to the right and level with the head of the subject. 3. At the beginning of an experiment the release device holds the platform at the base of the tower. The elastic ropes are not under tension. All preliminary preparations (mounting of camera, seating and securing of subject, balancing of platform, etc. ) are carried out under these conditions. After the elastic ropes are placed under tension, the platform is released and stopped only after the completion of the series of vertical oscilla- tions. 4. During the oscillations in the vertical plane, the subject is exposed to accelerations of varying magnitude. The values ob- tained during a typical experimental run on the subgravity tower are as follows: TABLE I. Time Sequence Duration of Subgravity (Seconds) Amount of G (Seconds) Stationary 1 0-0.75 1; 3; 1; o 0.75 - 2.35 0 1. 70 2.35 - 3.85 0; 1; 2.8; 1; 0 3.85 - 5. 15 0 1. 30 5. 15 - 6.65 0; 1- 2.5; 1; 0 6.65 0 1. 00 7. 65 braking action 16

Figure 5. Subgravity Tower The maximum distance covered by the subject is about 7. 70 m (from the moment of release to the highest point of the vertical trajectory). The force exerted through the elastic ropes in a typical experiment amounted to about 300 kg. One set of ex- perimental G-exposures lasts about 8 seconds. It was repeated 10 times for one experiment on physiological or psychological responses to subgravity. 5. The system can be designated as an "open loop" system because the subject has no control of the situation. Film coverage of the experiment and actual performance tests were carried out. 17

6. The instrumentation of the Subgravity Tower was rather crude; but a multichannel was installed for recording the EKG and heart rate. Other physiological and psychological measurements of the effects of changing accelerations and subgravity states of short duration can be taken by means of improved equipment. 7. Although the majority of test subjects displayed rather dismaying symptoms similar to those observed with seasickness, no serious trauma or injuries can be expected in experiments em- ploying accelerations and accelerative changes within the range given above. 8. Cost figures not available. 9. Information on projects and workload not available. 10. References are included in the general bibliography given at the end of this report. The Gravitron 1. The Gravitron was suggested by H. Walton as a device to study the subgravity conditions and the zero-G state at ground level. Since Mr. Walton died in an aircraft accident on Novem- ber 5, 1955, no action has been taken to construct the device; but the principle of his proposal is worth-while to be considered for realization. 2. The basic idea of the gravitron is to induce a state of free fall in a vertical tube. At the end of the fall the motion is reversed by means of deceleration-acceleration so that the same tube can be used for the next phase of subgravity. Figure 6 shows the I-tube gravitron. The subject enclosed in a pressure chamber falls freely in segment "a" of the vacuum type vertical tube. In segment "b" he decelerated to a standstill; then accelerated upward again to repeat the cycle. The schematic drawing in the right half of Figure 6 shows a different arrange- ment, one in which the only external power input necessary is the relatively small amount to compensate for friction of the cabin gliding wheels. The reversal in the direction of motion occurs in a curved part of the device. Care must be taken however that no jerks occur by the change of the directional vector component during the transition from vertical fall to non-linear motion. The deceleration due to friction in the curved section of the tube may be compensated for by added acceleration. 18

Ground Deceleration Acceleration I TUBE GRAVITRON U TUBE 6RAVITRON (From H. Walton. Jr.. Journ. of Aviat. Med. 28:291-294, 1957) Figure 6. Gravitron 3. Figure 7 shows typical velocity and gravity cycles in an I-gravitron. Zero-G states are alternating with states of in- creased acceleration. The vertical tube has a height of 350 m (1,148 feet). This would allow a zero-G state of about 14 seconds, and deceleration-acceleration at 4 g for 4. 66 seconds. Figure 7. TIME, SECONDS (From H. Walton. Jr.. Journ. of Aviat. Med. 28:291-294. 1957) Height, velocity and subjectively felt gravity force in an I-tube gravitron, of hypothetical dimensions: height of free fall "b" = 240 meters; height for deceleration-acceleration "a" = 1000 meters. Notice that subject is alternately in gravity free state for fourteen seconds and under 4g for 4. 66 seconds (lg = 980 cm sec -2). 19

Figure 8 shows typical velocity and acceleration cycles in an imaginary U-gravitron of 240 m (750 feet) height and 120 m (375 feet) radius. Fourteen seconds of zero-G or subgravity would alternate with 5. 6 seconds at 4 G. However, it is necessary to adjust design of the curved section to allow for a smooth onset of deceleration. 4. Table II gives the mathematical treatment of accelera- tion, velocity, and course in the device, neglecting frictional forces between cabin rollers and tracks. Looking at the accelera- tion pattern as given in Table II, we see that there is an abrupt change from the linear acceleration at the end of the straight section of the U tov at the beginning of the non-linear section. Thus, the value of r changes from infinite* to a finite one; and this results in an unsteady change of motion (change in accelera- tion) at a fraction of a second. A smooth transition from linear to radial acceleration must be provided if a gradual onset of the G-forces is to be obtained. 5. There is no control system particularly designed to operate in the gravitron. Certain principles of other free fall and acceleration devices could be applied and incorporated, how- ever. This pertains specifically to performance testing and scoring, telemetry, photography, and other data collecting tech- niques. 6. The device could be used to investigate some physical effects of short zero-G states; such as, convection, diffusion, dilution, and heat transfer of gasses and liquids; and equipment to be used in zero-G conditions, such as, electronic cooling equipment that relies on convective air movement or on boiling a liquid and recondensing the vapor within the system; ventila- tion devices, some uncooled electric lights; gyros, artificial horizons, bank indicators, rotometers and manometers, level indicators and controls; hydraulic heads, systems, drives, and bearings, lubrication systems; conventional purge boilers, dry sump lubrication of gear boxes, and artificial gravity devices (regenerative algae systems). 7. For maximum safety, the height of the device should not exceed the measures given in Figure 7, and the acting G- forces should be kept at or below the values given above, if no special damping equipment or anti-G suits are used. It is possible to adjust design of both I- and U-models to secure a *During the free fall in the straight part of the gravitron r = oo. 20

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gradual onset of G-forces. For further precautions see "WADC Vertical Deceleration Tower (Drop Tower)". 8. No cost figures or estimates available. 9. No projects formulated as yet. 10. References included in general bibliography at the end of this report. TIME. SECONDS (From H. Walton. Jr.. Jour n of Aviat. Med. 28:291-294. 1957) Figure 8. Height, velocity, and subjectively felt force of gravity a = in a U gravitron of hypothetical dimensions: 240 meters, "r" = 100 meters. Subject is alternatively in gravity free state for fourteen seconds, and under about 4. 7 g for 4. 3 seconds (lg = 980 cm sec-^). George C. Marshall Space Flight Center Vertical Linear Accelerator ("Pogo Stick") 1. The George C. Marshall Space Flight Center Vertical Linear Accelerator ("Pogo Stick") is located in Building 4487, Guidance and Control Laboratory, George C. Marshall Space Flight Center, Huntsville, Alabama. It was constructed in 1958 and has been modified several times since its first installation. The device is operated and maintained by a staff of four. 2. The Vertical Linear Accelerator is a pneumatic device which generates linear acceleration for test purposes. It consists of a carriage which moves freely up and down on rollers between a set of vertical guide rails, a cylinder located at the bottom 22

of the accelerator, and a rubber diaphragm which separates the cylinder from a large air tank. A picture of the device is in- cluded in Figure 9. The carriage, consisting of a test table, piston and its sup- porting framework, is set into motion by pushing the start button and manually adjusting a hand-operated valve until the desired amplitude has been reached. The carriage falls within the guide rails until the piston, 'beneath and integral with the table, enters the cylinder near the bottom of its travel. The work done in adiabatically compressing the trapped air reduces the kinetic energy of the falling piston until the pressure in the cylinder reaches the pressure in the air tank. At this time the diaphragm between the cylinder and the tank starts to move away from the cylinder side of the diaphragm housing. The kinetic energy of the piston has not been exhausted so it continues to move down- ward, displacing the air beneath it at constant pressure. It is this period of constant pressure that imposes constant decelera- tion on the piston, thereby producing the desired G loading. The G value obtained is a function of the tank pressure and the piston travel, and may be varied within the limits of the machine. After the lowest point of piston travel has been reached, the stored air in the diaphragm housing accelerates the piston upward with approximately the same G value reached during deceleration. Constant acceleration continues until the air in the diaphragm housing is completely expelled. Stored energy from compression continues to move the piston, but with decreasing acceleration, until it leaves the cylinder. Due to energy losses from friction and heat, the piston will not regain its original height. Therefore, it becomes necessary to introduce additional air into the cylinder. A cam operated con- tact switch, actuated by a cam rail on the carriage, operates a solenoid valve which feeds the make-up air into the system, thereby regulating the height of piston travel. The characteristics of the linear acceleration vs time can be modified to prepare the accelerator for specific tests on sub- gravity requirements. The maximum velocity obtainable with this device approximates 26. 6 ft/sec. The peak acceleration can be set continuously between 5G and 30G for payload weights up to 80 pounds. The capabilities and limitations of the device are given in Table III. By adjusting the deceleration pattern, states of sub- gravity and zero-G can be obtained. A special feature of the device is the capability of cyclic operation which is essential for 23

(US ABM A) Figure 9. Vertical Linear Accelerator the performance of long duration tests. Although the free fall is limited by friction of guide rails and air drag, a period of 1. 6 seconds of zero-G and subgravity can be produced at the rate of 30 cycles/min. 4. Theoretical basis and equations representing the sub- gravity conditions similar to the ones applying to free fall and vertical motion as given before. 5. The control system is manually operated. 6. The instrumentation consists of 22 shielded leads, air supply, and RF connector. 7. There are automatic safety devices which block the device from operating if the door on either the upper or lower level is open, if the hoist is not making contact with the switch at the top of the tower, or if the air pressure in the main tank is below 15 psi. If the piston comes within four inches of the safety block in the bottom of the cylinder, an alarm will sound; if it comes within 2 inches of the block, the power will auto- matically be cut off. There is a remote control stop which the operator may carry at all times when the accelerator is in operation. Moreover, the hoist meter will not operate if the crank for manually operating the hoist is not hanging on its hook. 8. Approximate cost of the device and its use not available at present. The person in charge of the operation is Mr. W. N. Allen, Applied Research Branch, Guidance and Control Labora- tory, Marshall Space Flight Center. In case of participation of 24

outside agencies, Dr. W. Haeusserman, Director of the Guidance and Control Laboratory should be contacted. 9. By means of this accelerator, highly sensitive devices (for example, gyros and drift rate indicators) are subjected to the influences of linear acceleration and subgravity, whereby their functioning is continuously controlled. The projects included ex- periments on air bearings, missile instrumentation packages, sea urchin eggs, and satellite instrumentation packages. The work- load during the previous year has been relatively small. The schedule for the next year is unknown. 10. "Operation Instructions for Vertical Linear Accelerator", prepared by the Development Design Section, Electro-Mechanical Engineering Branch, Guidance and Control Laboratory, and by Mr. J. Boehm, Guidance Design Section, Guidance and Control Laboratory, George C. Marshall Space Flight Center, Huntsville, Alabama. TABLE HI. VERTICAL LINEAR ACCELERATOR Capabilities and Limitations of Test Apparatus Linear Acceleration (InG) 30 20 15 Approximate Duration of Peak Acceleration (In m Sec) 0 0 50 220 Capacity Volume = cylinder; 10" diameter, 15" height Weight = 80 lbs. The Sulinac (Super Linear Accelerator) 1. The Sulinac is a device proposed by the Guidance and Control Laboratory, George C. Marshall Space Flight Center, Huntsville, Alabama. The construction of the device is pending upon receipt of available funds. 25

2. An artist's conception of the proposed device is given in Figure 10. 3. The limiting size of the payload will be about 23"x32"x44" (the latter dimension indicates the weight of the payload); the weight limitation will be about 600 pounds. The motion will be vertical linear; the acceleration and time spectrum is given in Table IV. The maximum velocity will be 64 ft/sec. The free-fall motion enhanced by air jets for compen- sating frictional losses will provide a duration of the subgravity state of about 4 sec/cycle. The operation of the device at 13 cycles/min can be maintained for any period of time. (US ABMA) Figure 10. Proposed Sulinac (Super Linear Accelerator) 26

TABLE IV. Sulinac Capabilities (Based on a Payload Weight of 200 lbs. ) Approximate Duration of No mear Peak Acceleration No. Acceleration (In G) 10 250 I 20 155 30 115 30 97 n 40 79 50 67 30 77 50 52 ra 70 41 90 33 100 30 CAPACITY: VOLUME = PRISM WITH MOUNTING BASE 19. 75"x31. 5"; HEIGHT = 44" WEIGHT = 700 LBS. 4. Theoretical basis of equations as given before. 5. The control system will be semi-automatic; the device will be started manually; operation of the device and control of the payload through open loop. 6. The instrumentation will include records of the accelera- tive forces and of subgravity and zero-G. Moreover, 120 signal leads will be available for recording technical and scientific in- formation, as indicated in ABMA research projects concerning the use of the Vertical Linear Accelerator. 7. The device will have completely interlocked safety features. 27

8. The estimated costs of the device are approximately $250,000, and of the housing and laboratory building also about $250,000. 9. Data not available. 10. Design and development was performed by the Guidance and Control Laboratory, ABMA. Dr. W. Haeusserman and Mr. J. Boehm can be contacted for further information. George C. Marshall Space Flight Center Space Flight Acceleration Simulator 1. The construction of a Space Flight Acceleration Simu- lator Facility, utilizing natural differences in altitude on earth (such as that to be found in the Grand Canyon and Yosemite Park, California), has been proposed. An artist's conception of the device is given in Figure 11. ^ ^ (US ABMA) Figure 11. Artist's Conception of ABMA Space Flight Acceleration Simulator Facility 28

2. It consists of track captured capsule traveling on cir- cular horizontal track capable of switching into a vertical track and return. Typical track, switch, and capsule details are shown in Figures 12, 13, 14, 15, and 16. There are no apparent limi- tations on payload size or weight. The sled considered herein is a 2-ton, 8-foot diameter, 24 feet long, turbo jet powered version carrying four men. LENGTH RADIUS VERTICAL TRACK 3622 — TRANSITION TRACK 1395 1 2 10 — 500 CIRCULAR TRACK 3142 500 TOTAL HEIGHT 4519 (US ABMA) Figure 12. Track Layout INSTRUMENTATION ^TURBOJET FUEL7 '•• ^Mflaeatij— * Ig / , \f \ \ \ ; 1 PRESSURE SYSTEM ^SWIVELLABLE EXHAUST (US ABM A) Figure 13. Typical Sled 29

INSTRUMENTATION PICKUP v. SPLASH SHIELD INTAKE EXHAUST (US ABM A) Figure 14. Typical Sled and Track Section TRANSITION TRACKS SWITCHING ARRANGEMENT > CIRCULAR POSITION SHOWN (US ABM A) Figure 15. Switching Arrangement (US ABMA) Figure 16. Artist's Conception of Sled Unit 30

ACCELERATION G M WEIGHTLESS PERIOD-- ORBITAL DURATION (US ABMA) Figure 17. Typical Acceleration Pattern Two-Stage Booster 3. Figure 17 shows typical acceleration-time history to be duplicated. Figure 18 shows simulator capability to duplicate the acceleration pattern giving sled (capsule) velocity and distance figures for the typical 500-foot radius track. Figure 19 indicates acceleration as a function of velocity and track radius. Figure 20 shows the height of the vertical track vs time at zero-G and com- pares time-around-the-clock and time at zero-G for given veloci- ties. Figure 21 indicates velocity buildup for longitudinal ac- celerations vs time during the travel on the circular track. Figure 22 covers the distance traveled by the sled and the deceleration forces involved for given velocities and time parameters. ACCELERATION (G) ebo S20 MO 260 tt0 300 TIME(SEC) lif0O LEGEND: . RESULTANT ACCELERATION CENTRIFUGAL ACCELERATION EARTH'S ACCELERATION LONGITUDINAL ACCELERATION OR DECELERATION TIME[SEC( 3QJSO 37394 51.910 DISTANCE TRAVELED (FT) (US ABMA) Figure 18. Typical Two-Stage Boosted Capsule In operation the vehicle starts on the circular track making one or more circuits as required, during which the vehicle is subjected to longitudinal, earth-G and centrifugal forces. The resultant of these forces is displayed in Figure 18 as running from 0 to 180 seconds and 210-300 seconds. Stage burnout at 90 seconds 31

vtLOornr (FT/MO VELOCITY IMI/HR) / s (US ABMA) Figure 19. Track Curvature Considerations VERTICAL TRACK HT/FT (US ABMA) Figure 20. Sled Velocity Parameters (US ABMA) Figure 21. Longitudinal Acceleration Study 32

(US ABMA) Figure 22. Sled Deceleration Considerations is simulated by proper programming of the continuous water brake which reduces velocity and therefore centrifugal force. The zero- G state is obtained during the vertical up and down "flight" after the switch to this track. Thrust is maintained to offset frictional and drag losses, thereby obtaining zero-G within sensor tolerances. 4. Simulation of subgravity other than zero-G is possible by entering the vertical motion with the variable vertical thrust com- ponent required. Control of the sled thrust vector is by a 180° swivel nozzle. The typical design shown allows a true zero-G capability to be maintained for 30 seconds. Figure 23 illustrates the increase in time possible for producing various levels of sub- gravity. / >,* ^^J^ ,-— ' — > 1 p-^-^ — «• ••^^^ / f / / ' 0 4 J 0 9 0 * 5 H } • 0 • D K C l> o a 0 TOTAL TIME OF GRAVITY LEVEL SIMULATION IMC.) (US ABMA) Figure 23. Endurance of Sub-Gravity Simulation for a 3622. 5 Ft. Track Other programs might include physical and biomedical ex- periments under sustained high G forces by refueling of the sled 33

or through electrical drive units. Ejection of manned or instru- mented capsules may be accomplished under acceleration or weightless states. Angular, oscillatory and jostling accelerations may also be added to provide any desired combination of forces. 5. Physical orientation of the subject within the sled is optional. Preselected or variable orientation is possible for any "flight". * Subject can be part of an open or closed loop. 6. All standard track type instrumentation systems are possible with this simulator. Due to the liberal sized cabin, either a board viewing of the subject or trainee or TV to the ad- joining laboratories is possible. Recording channels and noise levels are not known. 7. Since the device is ground-based and track-captured, the safety problems associated with free flight or controlled air- craft flight are eliminated. Other safety considerations are ob- vious for the circular track operation. Several systems exist for halting and holding the sled on the vertical track such as those used in elevator shafts, etc. 8. Facility cost estimated at approximately 20 million dollars, which would include the technical facilities, the technical support, administrative and logistical support facilities, housing and community facilities access and utilities. Operational costs would be relatively low due to the possibility of round-the-clock opera- tion of the sled itself. 9. The simulator would provide a highly desirable link in the selection and training of future astronauts. The ability to observe the candidates with an operator on board, to increase the performance outputs of the system, to allow the subject to be in a mockup of his flight vehicle, and to perform his actual duties during the "flight" should increase the number of acceptable space flight personnel. 10. ABMAReport No. DSP-TR-1-59 entitled "Space Flight Simulator", dated 16 March 1959. *Theoretical basis and equations representing the physical forces acting on the subject are similar to those given for the Gravitron. As in case of the latter, the characteristics of the ABMA Space Flight Acceleration Simulator does not provide for a smooth transition from linear acceleration into angular acceleration. In order to avoid jolts during the change from linear to angular ac- celeration and vice versa, another transition than proposed seems to be necessary. 34

B. Aircraft From the laws of kinematics, it can be followed that the true state of zero-G can be obtained in a dynamic situation only. The aircraft, with its six degrees of freedom of motion, seems to be the most promising means available at present for human experiments in the zero-G state. Some general remarks con- cerning the motion, acceleration and velocity characteristics will precede the discussion of particular types of aircraft used in zero-G research. Within the gravisphere of the earth, an aircraft is truly weightless when it is flying on a Keplerian trajectory. The char- acteristics of such a curve indicate that it can be represented by a parabola with vertical axis. Strictly speaking, the zero-G parabola is a trajectory described by an unpropelled body in ideally frictionless space subjected to a centrally symmetric gravitational field. Generally, such a trajectory is a conic, one focal point of which coincides with the center of attraction around which the body revolves. For sufficiently small velocities, such as are achieved with present day aircraft, the conic is a very elongated ellipse. The small section near the apex of the ellipse, emerging from the surface of the earth, can well be represented by a parabola. The condition for good approximation is that the dimensions of the section are small compared to the radius of the earth, or alternately, that the part of the earth's surface arched over by the section can be considered flat instead of spherical. For the mathematical analysis of a parabolic flight pattern, which is a vertical planar maneuver, an orthogonal coordinate system can be employed. The angle of climb is then determined by the direction of the vehicle at the beginning of the push-over; that is, at point 0 in Figure 24. The important information to be obtained from an analysis of the curve concerns the duration of the weightless state and its dependency upon other flight parameters. In ballistics, the duration of the trajectory is defined as the time required for the projectile to again reach the level of its initial projection. In our flight pattern, T is the duration of the weightless state. From Figure 24 it can be seen that the vertical location of any point of the parabola can be expressed by: y = vot sin f - g t2 (31) I 35

Figure 24. Schematic of Parabolic Arc. In the absence of gravity, a craft starting from point 0 at an angle t would follow the straight line vot sin t. Gravity causes the craft to fall. The vertical line segments between the straight line and the parabola represent the height of fall at the corresponding points of the trajectory. Also shown are directions and relative magnitudes of the velocity and its horizontal and vertical components. The velocity reaches its minimum at the peak P of the parabola where the vertical component vanishes. The magnitudes of the speed at beginning and end of the parabolic arc are equal. The same holds for the angle «. where vo = the initial velocity of the craft, and < = the angle of climb. In order to find T, we determine the value t for which y again becomes 0: T 2 L = _ v0 sm e . g (32) Equation (32) shows that T depends upon the value of g, the velocity vo, and the angle of climb t, the latter quantities measured at 0. At a given vo the longest duration is obtained when e = 90°. How- ever, the aircraft cannot rotate about 180° at zero speed on top of a straight-up ascent. Hence, our problem is to find the optimal 36

value of < withrespect to the initial velocity VQ and the minimum maneuvering speed of the airplane vmin. The optimal angle of climb can be found by considering the velocity diagram at point P in Figure 24. Since the horizontal component of the velocity must be constant, vx = vmm. Thus, we write: vx « vmin = vo cos* , (33) cos« = vmin . (34) Equation (34) shows that the optimal angle of climb depends upon the ratio of excess thrust and minimum controllability speed. Finally, we want to know what relationship exists between the duration of the weightless state and the peak altitude of the maneuver. The maximum height hmax is defined as the greatest vertical distance reached by the aircraft as measured from the ground. In Figure 24 the maximum height is that of the peak P of the parabola: "'max = hmin v° .2 2 sin* e : (35) 2g Another simple representation results, if the peak of the parabola is chosen as origin of the coordinate system. In this system, the altitude at any time is y= hmax -ft2, <36) the downward vertical component of the velocity is vy = gt. (37) From the total velocity __^^__^__ v = V vx2 + g2 t2 (38) the duration of one leg of the parabola is found 1=7 / vo2 - vx2 , (39) 37

where V voy = v - v <40> is the vertical component of the velocity at the endpoints of the parabola. In terms of maximal and minimal altitudes, this vertical component and the total duration can be represented by • V 2g<h voy • 2g<hmax - hmin> -r- ? g (43) Finally, the optimal angle of climb is again given by formula (34). In Figures 25, 26, and 27, the total duration T of the maneuver, the height hmax - hmin of the parabolic arc, and the optimal angle of climb «op^ are given in the range of maximal speeds between 400 and 2000 knots for three values of the minimal speed. Convair C-131B 1. The Convair C-131B Aerospace Medical Airborne Lab- oratory is located at WCTOC, Wright-Patter son Air Force Base, Ohio. The C-131B aircraft was built in 1953 and designed as an electronic test bed facility with modifications of increased flooring structure to handle electronic test apparatus. The Aerospace Medical Laboratory projects are monitored directly by two officers and one airman. 2. A photograph of the Convair-131B is attached (Figure 28). Weight limitations are programmed by aircraft station (approxi- mately 350 lbs/4 ft) and normally the volume requirement (33 1/2X 75"X26" test racks) is the limiting installation factor. The free- floating area (25'X5I5"X7I) can be used for large portable apparatus, weight distribution permitting. (Ref. T. O. IC-131A-1 for full payload weight, vs fuel limitations. ) 3. The attached C-131B parabola dynamics chart is re- printed for your convenience (Figure 29). The X-abcissa repre- sents degrees of pitch, the Y-ordinate represents maneuver entry speed (our maximum diving speed is 266 knots IAS), the upper right curves represent "positive" G Pulleys parameters (our 38

150 KNOTS 200 KNOTS 250 KNOTS 400 600 800 1000 1200 vmox- 1500 2000 KNOTS Figure 25. Duration T of parabolic maneuver as related to maxi- mal and minimal speed. The duration is approximately proportional to the maximal speed the craft can achieve. The minimal controllable speed has little influence unless it amounts to an appreciable fraction of the maximal speed. maximum G load is 2. 5 for structural safety limitations). The vertical lines represent zero-G duration in seconds, reading from pitch over point at right (normally +45°) to recovery point at left (normally -35 to 40°). 4. Continuous head-to-toe recordings have shown sustained parabolas averaging 0 + . 01 g for 14 seconds. Fore and aft g re- cordings have shown more variation (0 + . 05 g). The latter prob- lem exists because of the difficult throttle programming required to overcome the relatively slow changing aircraft momentum during initial acceleration. The maneuver itself is easy to fly; however, the aircraft flies through potentially dangerous pitch 39

200,000 FEET 150.000 100.000 50,000 400 600 800 WOO 1200 1500 •moi 2000 KNOTS Figure 26. Height of the parabolic arc vs maximal speed, for a minimal speed of 200 knots. The height increases approximately with the square of the maximal speed. For minimal speeds of 150 and 250 knots, the curve is shifted up or down, respectively, by approximately 1000 feet. attitudes and we have restricted the maneuver to experienced pilots (Figure 30). 5. Each experiment requires its own specific gear; how- ever, the following equipment has been installed to cover most projects: a. Heiland 24 channel photo-oscillograph. b. Head-to-toe and fore-and-aft oscillograph recorders. c. Altitude, airspeed, pitch and roll oscillograph recorders. d. Sequenced 16 mm camera coverage. e. Response switch tie-ins to oscillograph (psychomotor tasks, value judgements, etc.). 40

•max Figure 27. Optimal value of the initial angle of climb as related to maximal and minimal speed. The optimal angle increases with increasing maximal speed and approaches 90° for very large speeds. (Official USAF Photo) Figure 28. Convair C-131B Aerospace Medical Airborne Laboratory 41

5 - n it u it (Official USAF Photo) Figure 29. Velocity-Time-Acceleration Nomogram for C-131B Flight Parabolas (Official USAF Photo) Figure 30. Zero Gravity Experiments f. Physiological tie-ins to oscillograph (cardiac out- puts, respiration outputs, etc. ). Through experience with temperamental in-flight research it was learned to keep airborne gear as absolutely simple as possible. Mission aborts become very expensive and as a result the pro- grams require portable, simple-to-repair and reliable test equip- ment. 6. (Covered by Paragraph 5.) 7. The loss of oil pressure with inadvertent "negative G" operation has necessitated the addition of propeller pitch locks and a supplementary oil expulsion system. Fighter type battery caps were added to prevent acid spillage. The aircraft undergoes periodic structural inspection to foresee possible structure failures. Head helmets are normally worn by free-floating subjects over a mattressed area. 42

8. Aircraft use is requested through WCLDPPS, WADC, W-PAFB, Ohio. Formal arrangements are then signed with the Programming Office, Directorate of Flight and All-Weather Test- ing. Experiments in non-areo med experiments, such as fluid be- havior and tool evaluations under weightless conditions, are routed by WCLDPPS to the appropriate laboratory for direct control; however, these programs are still monitored by WCLDPPS. The only costs involved may be film and oscillograph papers procure- ment because the aircraft has been assigned full time to the Aerospace Medicral Laboratory for in-flight research. 9. Current physiological and psychological research efforts: Project Project Engineer Weightless effects on entering supine seat Captain Morrison Effects on cardiac output Dr. Zinn Effects on vestibular canal Dr. Dzendolet Effects on self-locomotion Mr. J. Frost Effects on non-torque tools Mr. Spencer Effects on pigeon orientation Dr. Belleville Effects on positioning mass Mr. Reese Effects on visual measures Major Pigg - Captain Pitt Effects on self-propulsion and muscle force Dr. Hertzberg The following experiments were prepared for the Aerospace Medical Convair during the period 1 July 1959 to 1 October 1959: No. No. Experiment Fits. Subjects Runs/Fit. Magnetic Devices - Channel 4 2 magnetic sandals and electro- magnetic shoes, "Project Hotfoot", are being completed for weightless walking (Simons, Nicholson). Semi-Supine Seat - Free-floating 2 2 subjects will enter and exit seat during weightlessness (Caton). 43

Experiment Propulsion Units - Velocities obtained by known thrusts will be studied with four sequenced cameras (Hertzberg). Self- control capabilities will be studied (Gardner). Project "Skyhook" - Rotational velocities will be measured (Simons, Hall). Non-Torque Tools - A special lock washer and non-torque wrench will be applied by a free-floater (Spencer). Psychomotor Panel - Duration measures of switching motions on six controls will be studied. Pitch, G's, and control actuation will be recorded (Wade). (Fig. 31-b) Vestibular Canal - Subjects will record rotation sensations as pitch, roll, and G's are recorded (Dzendolet). Animal Performance - Four pigeons will be released in various positions during weightlessness on two flights. Animal feeder and watering units will be tested (Sears, Grunzke). Mass Discrimination - Vertical and horizontal mass positioning of weights will be completed (Rees). Visual Acuity - Ortho-Reter trials on harnessed subjects will be run (Pigg, Kama). Pressure Suit - Astronauts will free-float with integrated oxygen supply units (Vail). No. No. Fits. Subjects Runs/Fit. 12 30 12 16 12 12 44

No. No. Experiment Fits. Subjects Runs/Fit. Self-Locomotion - Subjects will 4 2 generate four arm and leg move- ments in front of grid and free- floating mass (Frost). Flash Blindness Goggle and 3 1 Gradient Density Lenses will be subjectively evaluated (Alder). Retinal Deformation - Eikonometer 3 3 tests will be administered during weightlessness (Alder). Blood Pooling - Cardiac outputs, 3 3 pitch, and acceleration will be recorded during five tilt condi- tions (Olson, Zinn). Skin Resistance - GSR measures 4 8 will be taken on harnessed and unharnessed subjects during weightlessness (Johnson). Fatigue Flight - A subject will 1 1 fly a seven-hour cross-country mission in the copilot's seat with intermittent GSR recordings made (Ruff). Many of the above experiments will be flown concurrently. The Propulsion Laboratory (Fluid Behavior During Weightlessness) will use approximately 40% of the flight time and their short negative-zero G maneuver is not compatible with the tests. WADC ISO will fly news media personnel on several occasions. 10. Literature: Zero-G Experiments: Major E. L. Brown In-Flight Study of Magnetic Shoes: Captain J. C. Simons In-Flight Study of Stabilization Unit: Captain J. C. Simons Man's Ability to Apply Certain Torques While Weightless: E. Dzendolet and J. F. Rievley 45

A picture of three floating astronauts during a familiariza- tion flight in the C-131B is shown in Figure 31a. Figure 3la. Three of the NASA Astronauts free-floating in a C-131B during indoctrination and conditioning to the peculiar effects of weightlessness. Supplement 1. Walking Under Zero-Gravity Conditions The first complete experiment with magnetic shoes, which enable man to walk with an approximately normal gait under weightless conditions, was conducted by J. C. Simons (1959). A pair of magnetic sandals was made by attaching 12 Alnico magnets to aluminum soles. Although the plates were rigid, rubber shims allowed a *3° ankle rotation transverse to the walking surface. 46

Figure 31b. Psychomotor Performance Test During Zero-G in the C-131B. Each sandal required a break-away force of 21-22 lbs. at right angles to the contact surface. A plantar flexion of 8-10° was sufficient to induce an air gap between the magnets in the sandal and the floor to eliminate the magnetic force. A soft iron plate, 18"xl/8"xl3' was bolted to the ceiling of the rear part of the C-131B aircraft cabin. The plate served as a way for upside-down walking during the zero-G parabola (see Fig. S-1). This arrangement was chosen because of the need for reorientation of the subject. Results: The subjects had difficulty in walking with a normal gait. They walked flat-footed and stepped lightly, as though walking 47

on eggs. This was probably due to the broad contact area of the sandals (4. 5 inches square on the heel and 7. 5 inches square on the ball of the foot. ) No acute disorientation problems were found when subject entered the weightless condition from a supine position. Instead, all four subjects reported an immediate reorientation of "down" being where their feet were, as soon as their body rotation stopped. A basic index was formulated to define magnetic require- ments in terms of the inductive forces required to hold a subject stationary. Moreover, a vector analysis of the 1 G walking gait was made, which is also applicable to man's locomotion under zero-G conditions. Reference: Simons, J. C.: Walking Under Zero-Gravity Conditions. WADC TN 59-327, Wright Air Development Center, Air Research and Development Command, Wright-Patter son Air Force Base, Ohio, October 1959. (Official USAF Photo) Figure S-1. Walking with Magnetic Shoes Under Zero-G Conditions 48

Lockheed T-33A 1. One Lockheed T-33A each was assigned to the former School of Aviation Medicine, Randolph Air Force Base, Texas, and to the Aeromedical Field Laboratory, Holloman Air Force Base, New Mexico, in 1955. They were powered by a J-33A-35 jet engine developing 4,600 lbs. thrust. The aircraft is a tandem two-seater plane as shown in Figure 32. The pilot was seated in the front seat and the subject in the rear seat of the cockpit. A film camera (16 mm) was installed in the rear compartment for recording the subject's reaction to weightlessness and for related studies. Two pilots flew alternately, and a crew of four main- tained the aircraft. •urvemtu-ft; . Figure 32. Lockheed T-33A (Official USAF Photo) 2. The aircraft was not specifically equipped for use as a zero-G facility. In all the experiments flown, the test subject was accommodated in the back seat. In addition, small equipment for experiments and animals were carried along. 49

3. Subgravity and zero-G states were produced by flying aerodynamic parabolas as described before. A schematic picture of the T-33 zero-G flight pattern is shown in Figure 33. 21.000' (IT.OOO1) Figure 33. T-33 Zero-G SAM Flight Parabola: The maneuver starts at an altitude of about 18,000 feet, reaches almost 21,000 feet at the top of the parabola, and yields about 28 seconds of virtual weightlessness. Since the performance of this aircraft was rather limited, relatively high accelerations before and after the push-over maneuver had to be tolerated. In order to obtain subgravity states of about 25 - 30 seconds, a sharp dive was started at 20,000 feet with engine r. p. m. at 96%. As the IAS reached 350 knots, the aircraft was pulled up at about 17,500 feet at an angle of climb of 55°, producing a force of 3 G-units acting on the body. As the LAS dropped to 300 knots at full throttle, forward stick pressure was applied and a push-over initiated. At this point of transaction, a slight and momentary yawing often occurred, but was easily controlled by aileron motion. Slight changes in forward pressure of the stick were necessary to maintain the aircraft on the para- bolic arc. Minimum controllability speed was about 180 knots at a peak altitude of about 21,000 feet. As the apogee was reached, forward stick pressure was continued until the plane dived at 350 knots. When the angle of dive was about 55° again and the aircraft back at 17,500 feet, a pull-out was started early enough to prevent the aircraft from passing its safety limit and avoiding excessive ac- celerations during recovery. Usually at this point the maneuver was repeated using the pull-out acceleration for propelling the aircraft into the next parabola. 50

4. The theoretical basis and equations representing the physical conditions for the subject seated in the back seat of the aircraft has been given before. Since no special zero-G instru- mentation was available at that time, the accuracy of keeping the aircraft on an ideal Keplerian trajectory was dependent upon the skill and experience of the pilot. Since the latter was in control of starting and concluding the maneuver, he acted as experimenter insofar as timing and communicating with the subject was con- cerned. 5. The control system of the aircraft was purely manual, and the maneuver (including the amount of G forces acting on pilot and subject) was initiated and monitored by the pilot. An auto- matic device for producing and controlling the zero-G state had been proposed, but was not available for the T-33 zero-G studies. Thus, and because of safety features described below, only short periods (3-5 seconds) of zero-G were produced, and the remain- ing condition of subgravity was generally designated as "virtual weightlessness. " In addition, yaw and roll movements of the aircraft produced undesirable accelerations in all three axes be- cause of the sensitive aileron boost at low air speed, and the roll- ing motions exaggerated by tip fuel. 6. The standard accelerometer installed in all U. S. fighter type aircraft was employed as main reference. Further, the sub- gravity state was indicated by the floating of some small objects, and by the feel of lost support by pilot and subject. The scientific information was mostly recorded on tape and film and later evalua- ted on the ground. Experiments at that time mainly concerned the tolerance of human subjects to changing accelerations including weightlessness; studies of eye-hand coordination; the reflexes of cats and humans; and demonstration of the capability to react, eat and drink, and to demonstrate the weightless state pictorially. 7. Because of safety limitations, all experiments were per- formed while the subject was seated and held by the safety belt. Moreover, parachutes, 02-mask, and helmet had to be worn be- cause the flights were not without incident. During the subgravity state at altitude, fuel in the main tank had a tendency to evaporate, resulting in an overflow from the tank out through the vent drain line. This vaporization and lack of contact of the fuel prevented the pump from supplying enough fuel to the engine to maintain the required r. p. m. The few power reductions which occurred during zero-G maneuvers were corrected immediately by reducing throttle and producing accelerative forces. Complete loss of oil pressure also took place but was of little concern because the engine was 51

built to run for longer periods with no lubrication, hydraulic system was unaffected. The aircraft's 8. The cost for one flight, in which normally 8-10 parabolos were executed, was estimated in between $1,000 and $1,500. Al- though the T-33 is not an ideal aircraft for producing weightless- ness, it proved to be an economical and practical tool. It is still used at many USAF installations for training and other purposes. 9. No information available. 10. References on zero-G experimentation with the T-33 are included in the general bibliography at the end of this report. Lockheed F-94C, Zero-G Facility 1. The Lockheed F-94C "Starfire" is a two-seater inter- ceptor powered by a Pratt & Whitney J-48-P-7 jet engine, develop- ing with after-burner a maximum of 8,750 lbs. of thrust (Figure 34). For the sake of safety and lack of turbulence, the working altitude in between 20, 000 and 25, 000 ft. was found to be most satisfying. From this altitude a dive was entered, engine r. p. m. at 100%, and as the IAS reached 425 knots a climb of 65° to 70° (Official USAF Photo) Figure 34. Lockheed F-94C 52

from the horizontal was executed. With airspeed dropping off, forward stick pressure produced a stable and lengthy parabola yielding almost 40 sec. of subgravity and zero-G. 2. Since the F-94C has a fully pressurized fuel system, the power remained constant during the parabola. Stable in flight because of the 6° dihedral wing shape, the variable elevator boost made minor control changes easy and instantaneous, allowing small corrections to be made without varying the G-conditions appreciably. The high Mach rating eliminated the necessity of too early a dive recovery (Figure 35). 23.000' Figure 35. F-94C Zero-G "SAM" Flight Parabola: The maneuver is started at an altitude of about 18,000 feet, reaches about 23, 000 feet at the top of the parabola, and yields about 40 seconds of virtual weightlessness. 3. With after-burner, the longest time recorded was 43 sec. of subgravity (1956). In this maneuver at 20,000 ft. with elevator boost ratio at 11:1, engine r. p. m. was increased to 100% and the after-burner actuated. IAS was increased to 430 knots, and the aircraft eased up into a climb of 75° from the horizontal. The after-burner was used until this angle was reached to obtain an IAS of somewhat over 400 knots. At that time the after-burner was cut off and the parabola begun. The aircraft was pulled out at 12, 000 ft. at an IAS of somewhat in excess of 450 knots produc- ing a force of about 4 1/2 G for several seconds (Figure 36). 4. The theoretical basis and the equations representing the physical conditions of the subject were given before. In preparing the instrumentation, a sensitive visual indicator of the G-forces 53

involved was designed by Schock and Simons (1957) consisting of two microammeters, with a range of 25-0-25 micro-amperes, connected to a set of sensitive Statham accelerometers with a range of ±0. 5 G. One accelerometer was mounted to measure the forces acting in the vertical axis (z), and the other in the longi- tudinal (x) axis of the aircraft. Each division on the microam- meters was equivalent to ±0. 014 G. If the needles of both devices were on the zero mark, the pilot knew that he was exactly on the parabola (Figure 37). (Official USAF Photo) Figure 36. The F-94C Zero-G "Holloman" Flight Parabola: The maneuver is started at an altitude of about 20,000 feet, reaches about 25,000 feet at the top of the parabola, and yields about 45 seconds of virtual weightlessness. 5. Another set of sensitive accelerometers was fixed to the subject with a chest band to provide the recording to forces acting directly on the individual (Figure 38). Two Statham ac- celerometers were attached to a leather band which the subject wore around his chest. The output of these accelerometers was fed into a Century Model 409 recording oscillograph, a galvano- meter type instrument sensitive enough to record the output of the accelerometers without further amplification. 6. Pilot, subject, and experimenter were in continuous voice communication. The subject carried a tape recorder (Mohawk "Midgetape"); and the pilot could tune in the ground station whenever he deemed necessary. The subject was photographed during the maneuvers using a Bell & Howell B-1A camera with 54

(Official USAF Photo) Figure 37. F-94C instrumented cockpit. Visual zero-G display in front part of cockpit (left side of the picture) and base for mounting film camera in rear part of cockpit (right edge of picture). (Official USAF Photo) Figure 38. F-94C instrumentation. Preparation of test subject for flight experiment. On his left leg pocket is the Midgetape recorder; on chest belt is accelerometer for recording G-forces. 55

a wide-angle lens (Wollensack, F:l. 5, 89 degrees). The camera was mounted on top of the instrument panel in the back seat of the cockpit (Figure 39). (Official USAF Photo) Figure 39. F-94C zero-G instrumentation. Base of film camera, which was turnable and should be moved about 90° to the left, so that the subject would not be injured during ejection. The GSR - Dermo-Ohm-meter is installed below the camera base. For EKG and GSR recording an instrument rack was con- structed in the rear part of the cockpit, which contained an EKG amplifier (Grass, P-5), a "Dermohmeter" (Yellowspring-Fels, 22-A, airborne), the oscillograph mentioned above, and battery power supplies. Two EKG electrodes were placed on the chest; the other on the right ankle. The GSR electrodes were applied to the sole of each foot. All five electrode leads were combined to form one cable. A plug was provided for a quick release be- tween the subject and the amplifier, to avoid hazard or delay in case of emergency ejection (Figure 40). 7. Safety requirements identical to those applicable to ex- perimental flights in fighter type aircraft. 8. No cost figures available at this time. 9. No information available from the School of Aviation Medicine, USAF, Randolph Air Force Base, Texas, nor from the Aero Medical Field Laboratory, Holloman Air Force Base, New Mexico, where F-94C aircraft were employed in zero-G research. 56

•ECOflDINC OSCILLOGRAPH MIDWEST INSTRUMENT CO MODEL 560A (After H. von Beckh, Figure 40. Instrumentation for Zero-G Research in Fighter Aircraft F-94C. 10. References are included in the general bibliography at the end of this report. KG-135 1. A KC-135 tanker plane has already been modified for use as a "Zero-G Flying Laboratory. " Modification concerned (1) the three Sundstrant electrical oil drives to provide electrical power for the aircraft and test instrumentation under zero-G, (2) the hydraulic system and speed brakes to be used during the parabolic maneuver, (3) the engine oil system, (4) the fuel sys- tem to provide enough pressure to the pumps under conditions of weightlessness, and (5) provision for installation of test equip- ment and scientific instrumentation. The contractor is Boeing Airplane Company. 2. The characteristics of flight parabolas and of aircraft suitable for this purpose are similar to those of the C-131B; however, the large dimensions of the KC-135 will allow much more testing space (particularly with respect to unrestrained and free- floating subjects) and heavier test equipment of greater volume. A photograph of the airplane is shown in Figure 41. The gross weight will have to be limited to about 130, 000 pounds in order to keep within the safety requirements during the parabolas. 57

(Official USAF Photo) Figure 41. KC-135. 3. The performance date for the KC-135 aircraft has been calculated as follows: Weightlessness and subgravity time during one parabola 30 - 35 seconds Number of parabolas during one flight 20 Accumulated time of subgravity and weightlessness during one flight 10 - 12 minutes Maximum acceleration during pull- out before and after the parabola 3 - 3. 5 g Maximum angle of climb at start of parabola Working altitude Velocity at entry into parabola 50 degrees 25,000 - 35,000 feet . 8 Mach Calculations show that the CG will not shift enough throughout the parabola to affect the pilot's control over the vehicle. However, 58

it is anticipated that buffeting will occur at 25,000 feet, at the normal gross weight of 150,000 pounds, when the limiting Mach number is approached at a 2. 4 g pullout. Thus, less gross weight will have to be used as indicated above. 4. The equations and mathematical presentation of the phy- sical conditions were given above. Because of the larger weight and space available, practically any type of instrumentation for experiments with one or several subjects can be used. 5. The control system and instrumentation will be similar or advanced of that described under "C-131B, Zero-G Facility. " 6. Instrumentation and data reduction similar to that appli- cable for zero-G parabolas. 7. Safety requirements are same as that applicable for zero- G parabolas in aircraft. Special safety measures necessary in case of experimentation with rocket propellants and liquid H2 as required for the CENTAUR start tests. 8. No cost figures available at present. Since the first modified KC-135 will be stationed at Wright-Patter son Air Force Base, WADC personnel will monitor the test program. Requests for participation in this program should be directed to Commander, Wright Air Development Center, Dayton, Ohio, ATTN: Mr. H. R. Shows, WADC-WCLOR. 9. The research program for the KC-135 is actually the expanded C-131B research program. Information on the research subjects are listed below. a. Aerospace Medical Laboratory, WADC. Human muscle and body control, mass discrimination, microswitch actuation, non-torque tools, self-locomotion, human stress, eye protective equipment, supine seats, visual perception, human engineering for space, human performance in pressure suits, fluid action, heat exchangers, boiler condenser unit, airsickness and nausea, bonded magnetic materials, pigeon orientation, vesti- bular canal, magnetic suspension devices, visual activity, evalua- tion of oil expulsion, gyro-stabilized man, human propulsion unit, and animal performance for behavior. b. School of Aviation Medicine, USAF. Cardiovascular functions, blood circulation, pulmonary functions, and fluoroscopy and x-rays of the body and selected organs. 59

c. Jet Propulsion Laboratory. Venting and phase separa- tion of cryogenics, heat transfer of cryogenics, ion propulsion, capillary action tests, centrifugal orientation, mercury condenser, liquid N2 and O2 converters, space vehicle reaction control sys- tems, propulsion systems for human locomotion, energy conversion units, fluid motion effects, and CENTAUR restart tests. d. Aeronautical Accessories Laboratory and AEC. Deter- mination of heat transfer coefficients, and behavior of liquids and gases in pumps and turbines; experiments concerning nuclear power conversion loop using the Rankine cycle fluids may be sodium, potassium, mercury, sulphur, and rabidium; experiments on the effect of weightlessness in conjunction with SNAP I and II programs. e. Convair Astronautics. Experiments concerning the evaluation of CENTAUR's restart capability under weightless conditions. It is planned to install a 30-foot test vehicle con- taining a test cell 31 inches long by 12 inches wide. The cell will be suspended at three degrees of freedom so that pitching and yawing of the airplane will not affect it during weightlessness. 10. "Trip Report" by James W. Carter, Chief, Crew Engi- neering Unit, Missile System Analysis Section, Future Projects Design Branch, Structures and Mechanics Laboratory, Develop- ment Operations Division, Army Ballistic Missile Agency, Red- stone Arsenal, Alabama, dated 3 November 1959. TF-100F Aircraft 1. The School of Aviation Medicine, USAF, has two 1956 TF-100F aircraft that have been completely instrumented to con- duct weightless studies. These aircraft are maintained at SAM Flight Operations, Kelly AFB, Texas (Figure 42). The total staff and personnel assigned and engaged in our zero-G studies is fifteen. 2. The TF-100F aircraft has undergone extensive testing of various system components; but no modifications of the basic structure have been necessary. The enclosed photographs show the instrumentation that has been accomplished to record physio- logical response and performance data during flight. The size of the rear cockpit of the TF-100F aircraft limits the size of test objects to relatively small items that can be easily controlled by the subject and not jeopardize flying safety. 60

(Official USAF Photo) Figure 42. Lockheed F-100 3. Sustained periods of zero-G up to 60 seconds can be produced by flying the TF-100F aircraft through the parabolic flight maneuvers. The parabola is entered at a true air speed of about 655 knots following an initial dive and pull-out at 2. 5 - 3. 0 g. The minimum velocity at the apex of the parabola will be about 100 to 150 knots IAS (indicated air speed). The altitude covered from apex of the parabola to the final pull-out is about 40,000 to 20,000 feet. 4. As given before. 5. Control systems are standard aircraft control. Subject is not usually in the control loop since the TF-100F is not used primarily as a reaction control simulator, but subject can be in control loop if desired. 6. The instrumentation seen in Figures 43, 44, and 45 will allow 14 different parameters of information to be recorded con- tinuously and simultaneously. This can be recorded on the on- board oscillograph at the present time. The capability of tele- metering this information or recording it on magnetic tape is expected soon. 7. Normal safety regulations are followed. All connections to the subject are of the quick disconnect type. 61

(Official USAF Photo) Figure 43. F-100F Instrument Panel for Zero-G Tests (Official USAF Photo) Figure 44. TF-100F: Zero-G Test Instrumentation 8. Requests for use of these zero-G research planes should be made through official channels to the Commandant, School of Aviation Medicine, Brooks Air Force Base, San Antonio, Texas. 62

(Official USAF Photo) Figure 45. TF-100F: Connector Plugs for Zero-G Experimentation 9. The year 1958 was spent in instrumenting and testing of the two TF-100F airframes and engines. The final testing was completed in July 1959. The data recording systems have under- gone extensive development and modification since that time, and the coming years will see continuous use of these aircraft in carrying out the zero-G research program of the School of Avia- tion Medicine, USAF, Brooks Air Force Base, San Antonio, Texas. 10. References are included in the general bibliography at the end of this report. F-104A "Starfighter" 1. The Lockheed F-104A "Starfighter" is available in two versions: the fighter type aircraft for operational use (F-104A), and the two-seater version which has been designed for training purposes, and which can also be used for zero-G experimental flights (F-104B). It is available at several USAF fighter com- mands and also at the Test Flight Center, ARDC, Edwards Air Force Base, California. A picture of the F-104A is given in Figure 46. 2. Since the F-104A is one of the foremost operational air- craft, most of the performance data are classified. Specific data on speed, peak altitude, angle of climb, range, etc. can be found in the appropriate technical manual. From the data available, 63

it can be concluded that one subject can be carried along in the two-seater version, with the pilot acting as experimenter, similar to the case described before. (Official USAF Photo) Figure 46. Lockheed F-104A Starfighter 3. In this context, a special purpose version of the F-104 is described. An H^C^ jet reaction control system has been in- stalled in an F-104 aircraft to investigate roll, pitch, and yaw control parameters during high altitude flights at low aerodynamic pressure. The controls are actuated by a 3-axes, side-located controller which operates electric servo-valves. A photograph showing details at the wing-tip reaction control installation is included (Figure 47). 4. The theoretical basis and equations representing the use of the 104A as a zero-G facility can be deducted from the math- ematical treatment of flight parabolas given in the introductory section of the report. The characteristics for obtaining maximum periods of weightlessness are given in Table V. o 5. The rockets provide angular accelerations of 10°/sec in roll, and 5°/sec^ in pitch and yaw. The present system is open loop, but it is planned to incorporate a closed loop system (velocity command and attitude command) at a later date. 6. Instrumentation and channels for technical data and in- formation on weightlessness will be available. No details are available as yet. 64

(0fficlal USAF Photo) Figure 47. F-104 Reaction Control Instrumentation TABLE V Characteristics of Flight Trajectory of Various Types of Aircraft to Produce States of Weightlessness of Maximum Duration* nmax ~nmin (Feet) T (Sec) Aircraft (Knots) (Deg) C-131B 250 2,000 35 15 T-33A 350 5,000 55 30 F-94C 425 8,000 65 40 KC-135 500 10,000 50 35 F-100F 685 20,000 75 60 F-104A 800 30,000 75 1/2 80 X-15 4,500 500,000 _ _ _ 300 *The figures given in this table are approximations only. 65

7. Proposals to use the F-104B for the zero-G maneuver were not favorably received, however, because of anticipated lubrication problems of the engine similar to those observed in the F-100. They require lengthy and costly engine modification. The high altitude whieh will be reached in flying trajectories of a long zero-G duration will require pilot and subject to wear high altitude pressure suits. 8. Approximate cost of engine modification is $100,000 per aircraft. The cost of one flight is unknown. 9. Several zero-G flight experiments were made; the maxi- • mum duration of zero-G exposure was about 50 seconds. Installa- tion of the reaction control system was completed in 1959 and flight tests were completed which yielded information on jet reaction control during normal -gravity and in subgravity. 10. No information available. The X-15 Research Aircraft 1. The X-15 Experimental Research Aircraft will be avail- able for experiments at the NASA Flight Research Center, Edwards, California, in 1962. Flights for the express purpose of obtaining zero-G data are not anticipated in the near future, although certain flight profiles will impose some period of zero-G on the pilot. 2. Flight duration of 200 seconds at low dynamic pressure will be obtained. Moreover, these flights will provide appreciable test duration at zero-G, with the maximum duration depending upon the maximum altitude capability. (See Figure 48. ) It ap- pears that test durations of 300 sec. at zero-G may be feasible. However, since the airplane is still in the early flight test stage, the maximum zero-G capability will not be attained for some time. 3. The instrumentation for the X-15 flights will include the recording of accelerations in three axes, so that the duration of zero-G will be obtainable. Furthermore, it is anticipated that, incidental to these flights, some physiological data, EKG, body temperature, respiration, as well as subjective observations of the pilot, will be available. 4. An X-15 pilot training program was conducted on the centrifuge at the Naval Air Development Center, Johnsville, Pennsylvania, in 1960. This program covered the X-15 flight profile, particularly the occurrence of high accelerations. 66

(Official USAF Photo) Figure 48. X-15 During Powered Flight 5. The X-15 will carry over 1300 lbs. of instruments. The full-pressure suits to be worn were specifically designed with 24 electrical contact points for connections between the physiological sensors and the telemetry transmitters. 6. The physiological package to implement the measure- ments indicated above has been developed by North American Avia- tion for both the on-board recording and telemetering of physio- logical data. This system has a growth potential for recording additional variables, particularly those associated with and typical for zero-G effects. Evaluation of the North American Aviation physiological data package began at the Air Force Flight Test Center in March of 1959. 7. Safety features include the MC-2 full-pressure suit and the open ejection seat, incorporating fins and twin booms for roll and yaw stability. A 24-foot, back-type parachute is mounted in a special container on the ejection seat. Upon actuation of the ejection handles, the canopy deploys ballistic ally, initiating the proper sequence of events. Man-seat separation and parachute deployment are dependent upon present altitude and dynamic pressure sensors. The escape system will operate in the region of 600 knots IAS, or a dynamic pressure of 1500 lb/ft2. 8. No information available. 67

9. Projects as described above. 10. References included in the general bibliography at the end of this report. C. Ballistic Missiles As the jet aircraft is the best available means of studying the effects of subgravity and zero-G upon men, so the ballistic missile is the best means for investigating the effects of sub- gravity and zero-G on other biological specimens, including ani- mals, today. Ballistic missiles are generally accelerated during the initial part of their trajectory only; namely, from the start (or liftoff) to burnout (or cutoff). They are started vertically, be- cause this simplifies the launching operations and shortens the time necessary to penetrate the atmosphere. After this initial vertical climb the vehicle undergoes a programmed turn toward the target. During this turn the guidance system begins to func- tion and continues to do so until the desired altitude h, speed V, and angle 9 are attained (at B, Figure 49), whereupon it gives the FREE FLIGHT (Air Univ. Qtrly. Rev.. Vol. 9. Summer 1957) Figure 49. Trajectory of an ICBM signal for cutoff of the propulsive power. Perception and cor- rection of vehicle attitude, exercised by the control system, are continuous during the powered flight. Both the attitude of the vehicle and the motion of its center of gravity relative to the re- quired trajectory are adjusted by altering the direction of the thrust of the rocket engines, for instance, by putting jet vanes in the exhaust stream or by gimbaling the rocket thrust chambers. 68

There are many sets of values of the speed V, angle 9, and spatial position of B that will put the nose cone on a trajectory terminating at the desired target; but some sets are more favor- able than others in respect to amount of propellant consumed by the engines or required precision of aim. It is the function of powered flight to impart to the nose cone, as accurately as possi- ble, a favorable set of these parameters. The energy expended in propelling the vehicle during the powered flight increases with the weight of the vehicle. Because both the kinetic and the potential energies are approximately pro- portional to the weight of the vehicle at thrust cutoff, it is desirable that this weight be as little as possible in excess of the weight of the nose cone. This objective is aided very materially by divid- ing the vehicle into two or more parts, or stages, with each stage containing a rocket propulsion system. Launching is accomplished by starting the engines of the first stage and, in some designs, also of other stages. At some time during the powered flight the first-stage engines are shut down, and this stage is jettisoned from the remainder of the vehicle. The engines of the next stage are then started, if they are not already operating, and they propel the vehicle on toward B. As the missile nears B, the engines on the last stage are shut down, and the final adjustment of the velocity needed to keep the nose cone on a trajectory that will reach the target is accomplished with rocket engines of compara- tively small thrust, called vernier engines. Thus the term thrust- cutoff point B refers, accurately speaking, to the point where the vernier engines are shut down rather than to the shut-down point of the engines of the final stage. The trajectory beyond the thrust-cutoff point B may be divided into two segments: the free-flight portion, from B to the point C of re-entry into the atmosphere; the re-entry portion, from C to the impact point T (see Figure 49). For a long-range missile the free-flight portion BC is above the sensible atmos- phere; hence the missile during this phase is a freely falling body, the only force action on it being gravitational attraction. During re-entry of a freely falling body, the aerodynamic forces also come into play, and these slow the missile and cause it to become heated. The length and shape of the free-flight trajectory are deter- mined by the speed V of the missile at thrust cutoff, the angle 9 between the local vertical at B and the direction of V, the altitude h of B, and the values of the acceleration due to gravity g along the trajectory. 69

Considering a given point B and a given target T, one finds that for every thrust-cutoff speed V between the lowest and the highest values needed to reach the target, there are two values of the angle 9 that yield trajectories connecting B and T. One of these trajectories is steep, of high apogee; the other is flat, of low apogee. As one decreases the thrust-cutoff speed V, these two possible trajectories approach each other, the steeper tra- jectory becoming flatter, and the flatter trajectory more arched. In the limit, when V attains the minimum value for which the missile will reach the target, the two trajectories merge into a single one of medium height (Figure 50). Because this medium trajectory requires the smallest speed V, and therefore minimum kinetic energy at thrust cutoff, it is optimum with respect to pro- pellant requirements. It is also more favorable in other respects. For the steeper trajectory the re-entry speed is higher, thus pre- senting a more formidable heating problem. For the flatter tra- jectory the re-entry path through the atmosphere is longer. Both very steep and very flat trajectories require a more precise guidance system. Figure 50. Schematic Representation of Missile Trajectories with Different Launching Angles A simple picture of a free-flight trajectory may be obtained by considering first the case where the range and time of flight are so small that the missile can be assumed to be traveling over a flat and motionless earth, above which the acceleration due to gravity g is at every point the same in magnitude and always directed normal to the flat surface (Figure 51). For this situation, the horizontal range from thrust cutoff to impact can be obtained as follows: y = f t2 - vt cos 9 (43) 2 x = vt sin 9. (44) 70

(An Univ. Qlrly. Rev.. Vol 9, Summer 1957) Figure 51. Effects of the Earth Spin and Curvature on Missile Trajectory At impact time 1 " = f li - vt; cos9 q = vti sin 9 , (45) (46) (47) ti ^ = + v cos 9 2 t (48) vsin9 v sin 9 xi = 2v2 sin2 9 h 2v2 - + - sin 9 cos 9. (49) (50) Since we obtain h = xi tan (51) ^ = x + A x = 2v sin 9 (cos 9 + sin 9 tan 0), (52) g where Ax is the additional range gained because thrust cutoff oc- curs at B instead of on the ground at O, and where 0 is the angle between the horizontal and the straight line drawn from B to the 71

point of impact. As the range is increased, the effects of the earth's curvature and rotation become more and more important. Figure 50 shows the situation on an imaginary point in space. If the earth would be at rest, a missile leaving the thrust-cutoff point B would follow the same path as in Figure 51, except that OX is now to be regarded as the tangent to the equator at O. Let us see how this path is changed because the earth actually is ro- tating and its surface is not flat. In Figure 52 the coordinate system XOY is to be thought of as fixed in space, as not partici- pating in the earth's motions. This means that the origin O does not move and that the missile leaves B at the moment when B is vertically above O. (Air Univ. Qtrly. Rev.. Vol. 9. Summer 1957) Figure 52. Increase in Range Due to Earth Curvature and Rotation The trajectory is extended from D to F because the hori- zontal component of the missile's velocity at B is increased from the locally imparted value V sin 9 to V sin 0+ o»R, where w is the earth's angular speed of rotation and R is the earth's radius. The circumferential speed wR, which the missile has before it is launched and retains during flight, is about 1600/ft sec east- ward; this is a sizable correction even for missiles for which V sin 9 might be as much as 20,000 ft/sec. Note that for westbound missiles, this effect of the earth's rotation would reduce the length of the trajectory. For motion along any parallel of latitude /8 other than the equator, the correction would, of course, have the smaller value wR cos 0 eastward. While the missile is traveling from B to F, the point on the earth's surface directly beneath B has advanced from O to O'. This extends the trajectory to the point G because the impact area 72

has been displaced downward, from OF to O'G, during the missile flight. Such an extension would also occur for a westbound missile. At O1 the apparent horizon is the line O'H, which cuts the trajectory at H, and thus the trajectory is extended to H. Notice that this particular extension results from a downward rotation or tilting of the apparent impact area with respect to OX during flight. For a westbound missile the rotation of the impact area, as ob- served from O1, would be upward, resulting in a reduction of tra- jectory length. The trajectory is still farther extended, from H to 1, be- cause of the curvature of the earth, which gives the missile addi- tional time to acquire range. This extension is positive, no matter in what direction the missile is traveling, and would occur even if the earth were not rotating. The longer the range, the greater will be this extension, because the separation of the spherical sur- face from the plane OX occurs at an increasing rate as the distance from O increases. The missile would reach point 1 only if the gravitational force on it were at every point parallel to the Y-axis. Actually this force is directed toward the center of the earth at every in- stant of the flight. Consequently a backward component of gravi- tational force sets in as soon as the missile leaves the thrust- cutoff point B, and its magnitude increases steadily with the time since the missile left B. The net effect is to shorten the trajectory, so that impact occurs at some point J, rather than at 1. Actually the backward component of the gravitational force is associated with two different factors. One is the displacement of the missile from the fixed point O as a result of its locally imparted velocity V. This part of the backward component increases with the dura- tion of flight, decreases as the distance of the missile from the center of the earth increases, and would exist even if the earth were not rotating. The other factor is the departure of the missile from O because of its velocity wR resulting from the earth's rota- tion. This part of the net backward component is always westward, thus reducing eastward ranges and extending westward ranges. Although our interest has been mainly to show in a qualitative way how the rotation and curvature of the earth affect the range, it should be said that the method used here can be generalized to cover the case of a missile projected at any latitude and in a tra- jectory the plane of which is directed in any desired azimuth. For any case, however, the approximations involved in deriving the mathematical expressions for the various independent correction, 73

or perturbation, terms are least objectionable for missiles having small velocities at thrust cutoff. In order to compute trajectories in a simplified manner, we assume that the earth is a homogenous sphere and therefore at- tracts a missile as if all the earth's mass M were concentrated at its center (Figure 53). We have then a two-particle problem. (Air Univ. Qtrly. Rev.. Vol. 9, Summer 1957) Figure 53. Variation of Trajectory Due to Changes in Velocity that of a missile of relatively small mass m in free flight under the gravitational attraction of another particle, the earth, of ex- ceedingly large mass M. Notice that the only role played here by the earth's surface is to provide launching and impact areas for the missile. The trajectories to be used in coordinating the preliminary designs of the major subsystems of any particular type of missile are called reference trajectories. For this preliminary phase the trajectories will be sufficiently accurate if computed with respect to a nonrotating spherical earth. Thus the earth in Figure 53 is to be thought of as motionless in an inertial frame of reference-- a nonrotating set of coordinates in space that, for all present pur- poses, may be regarded as having its origin fixed with respect to the center of the sun. Newton's equations of motion then apply in 74

their simplest form, and from them an equation for the various possible free-flight trajectories of a missile may be derived. This equation turns out to be the general equation of a conic section. As to whether any particular trajectory will be a para- bola or an ellipse is found to depend on whether the ratio of the missile's kinetic energy to its potential energy at thrust cutoff is equal to unity or is less than unity. Knowing this, one can then show that the speed V of the missile at cutoff determines the type of path as follows: An ellipse with its farther focus at the earth's center if (53) \ F R + h where go is the Newtonian constant of gravitation, M and R are the mass and the radius of the earth, respectively, and h is the altitude at the cutoff points; a circle around the earth if v • V — ' <54> f R + h i. e. , approximately 5 mi/sec, and 9 = 90° (for other values of 9 the path will be elliptical); an ellipse with its nearer focus at the earth's center if VgoM / - >V<1/ R+hf 2goM . - ' (55) R+h i. e. , if v is between about 5 and 7 mi/sec; and a parabola or a hyperbola, if J I R (56) At these velocities and any value of 9 the missile will escape from the gravitational pull of the earth. Other factors, such as motion and gravitational anomalies of the earth, cannot be discussed in this context. 75

The principal characteristics of a missile used in bio- astronautic research is its ability to carry instruments and speci- mens to greater heights and for longer durations than will any other known type vehicle. However, the experimenter must keep in mind the dynamic characteristics of a rocket flight pattern, which will determine the type of experiment to be made and also affect the kind of measurements obtained. In experiments with a squirrel monkey in a U. S. Army JUPITER IRBM nose cone in December 1958, a subgravity period at about 0. 03 G was main- tained for approximately 500 seconds. Already, 13 months earlier, the first biosatellite had been placed in orbit, permitting the study of the physiological effects of zero-G on Laika during an exposure of about 8 days. Thus, technical and scientific space research, including zero-G by means of rockets and satellites, has already started and has yielded valuable data. Some of the advantages of this type of research tool are clearly defined by its flight characteristics, namely, the extremely high altitudes obtainable by means of rockets, and its powered flight outside and independent of any atmosphere. Its outstanding feature, of course, is that of producing zero-G for practically any length of time. This is naturally dependent on the development of rocket boosters powerful enough to ac- celerate a payload to reach orbital velocity. At present, there exists no rocket in this country that has been designed particularly for biological experimentation. Even the vehicles used in the "Discoverer" program of the USAF, are conventional IRBM and ICBM components. Moreover, this ex- perimentation is done almost exclusively on a non-interference basis with technological and military requirements. On the other hand, special sounding rockets have been particularly designed to meet the requirements of astronautical research. This may be one of the reasons why progress in space medicine is so slow and "falling behind engineering advances. " The first rockets employed for biomedical research in this country were the Aerobee (Figure 54) and the V-2. In both vehicles, weight and space for biological specimens and instrumentation were rather limited. For example, the nominal payload for the Aerobee was 150 lbs. , with which the normal peak altitude was about 70 miles. Space in an Aerobee was about 6. 25 cu. ft. In the large rockets, like the V-2, REDSTONE, JUPITER, THOR, ATLAS, etc. , much more space exists. Even so, the nose cones which usually carry the payload are crowded with equipment, mostly because of the need to obtain as many measurements as possible, 76

FIN PROPELLANT CHAMBER NOZZLE PARACHUTE BEACON - TUFTS RECORDER TELEMETER DYNAMOTOR ANTENNA THRUST STRUCTURE TAIL STRUCTURE -ANIMAL CONTAINER - BATTERIES (USAF Sch. of Aviat. Med.. Randolph AFB. Tex.) Figure 54. Aerobee Rocket and also because of the non-interference type of mission mentioned above. Some of the restrictions and limitations of this type of re- search must be pointed out to the rocket enthusiast, however. First of all, the great expense of rockets will doubtless often pre- vent the execution of an experiment despite its desirability. Most of the present-day rockets are in the million dollar bracket. Second- ly, they are still not reliable enough to guarantee the accomplish- ment of the mission. Thirdly, only a few of them are recoverable and special provisions must be made for the telemetering of bio- logical data, loss of the specimens involved, or their recovery by means of specially designed retrieving techniques. Fourthly, the payloads must be prepared so that they can withstand the long checkout periods on the pad before liftoff, the high accelerations and decelerations during the flight, and the stresses involved in re-entry and recovery. Generally, the nose cone is pressurized and climatized, but the bio-package must be protected against the hazards of space factors and secured in case of equipment failures. Fifth, most of the rockets and satellites available today are spun for stabilization. When it is essential to keep the specimen in a state of subgravity or zero-G, special precautions must be under- taken to stabilize the missile by gyros or to place the payload in the center of gravity or in the spin axis of the nose cone. Finally, 77

an IRBM or ICBM covers the distances of about 1,500 to 5,000 miles, during a 15- or 45-minute period, respectively, of which only a certain part is spent in zero-G (4 and 6 minutes in JUPITER flights AM-13 and AM-18, respectively; 40 minutes in the THOR- ABLE combination employed in Project MIA). The yield in zero-G duration is spectacularly greater through the use of larger boosters. At any rate, the significance of these relatively short times for the investment of large sums of money, long preparation time, and in- strumentation must be plain to prospective experimenters. Since we cannot duplicate the zero-G condition in ground- based laboratories, we are forced to venture into rocket flight experiments in weightlessness research. As in the past, we are faced with the necessity to establish the basis for experiments with men through the exposure of animals to new and unexplored conditions. Contrary to the subject in a manned vehicle experi- ment, the animal can be expended; and in this case, it must be instrumented and connected to a multi-channel transmitter. In the worst case we are faced with a test, the only yield of which consists of the experience gained through the preparations of the experiment. However, this and the results already obtained in rockets and satellites justify animal experiments on the biological and medical effects of zero-G as an essential step in achieving manned orbital and space flight. It does not seem to be coinci- dental either that the Soviet scientists have successfully carried out a larger number of space biological experiments in the last decade in conjunction with their development of larger boosters than was done in this country; and only the Sputniks could make possible a biological experiment, which concerned the effect of zero-G over a longer period of time. U.S. Army JUPITER (SM-78) 1. The U. S. Army JUPITER IRBM is an operational weapon system available at various missile installations of the USAF (Figure 55). Developed by the Army Ballistic Missile Agency, Huntsville, Alabama, it was used as carrier vehicle for biological payloads in 1958-59. The experiments were done by scientists at ABMA, the U. S. Army Medical R & D Command, and the U. S. Naval School of Aviation Medicine. The missile is now built by the Chrysler Corporation. 2. The missile signified the ultimate in simplicity. Its constant diameter cylindrical body is slightly under 60 feet in length and 9 feet in diameter. The control and guidance section 78

(US ABMA) Figure 55. U. S. Army JUPITER Missile is located in the conical instrument compartment on top of the liquid propellant tank. The nose cone, mounted on top of the in- strument compartment, is a continuation of the conical section with a bluff re-entry tip. Operational nose cones were success- fully recovered after 1,500-mile flights. Biological payloads up to 250 pounds (including life support system) were flown and re- covered. 3. JUPITER is powered by an S-3D Rocketdyne engine, developing approximately 150,000 pounds thrust. The motor is gimbaled to correct deviations from course. The body has no wings or fins. Roll control is provided by swiveling the exhaust of the gas generator. Zero-G and subgravity occur after cutoff and spin-up, respectively. Maximum acceleration during launch is about 15 g; maximum deceleration during re-entry about 38 g for several seconds. The range of the missile is about 1,000 miles, maximum altitude 300 miles, and velocity about 10,000 m. p. h. 4. The accelerations prevailing during free-floating have been calculated as practically zero; after spin-up, as varying 79

degrees of subgravity, lasting for several minutes depending on the beginning of spin-up. The location of the payload within the nose cone and its distance from the center of gravity as well as spin rate and stabilization determine the accelerative forces act- ing on the specimen. 5. The JUPITER IRBM flight control system is fully auto- matic. It does not allow for inflight monitoring of the payload. In some biological experiments an open loop system has been em- ployed to trigger the start and certain phases of the experiment. Since the missile control system is to function independently of payload and experimental design, separate systems must be used for payload control. 6. Number of recording channels depends upon the tele- metry system employed. Twelve commutated and 7 straight channels were available in U. S. Army Bio-Flight No. 2. Details of the experimental setup and measurements in Bio-Flight No. 1 are given in Tables VI and VII, and in Figures 56, 57, 58, 59, and 60. Telemetry measurements available in Bio-Flight No. 2 are listed in Table VIII. 7. Preparation and execution of the experiment must concur with range safety requirements for ground and flight operations. Payload will be destroyed with missile if deemed necessary by range safety officer. Figure 56. Oscillographs of Bio-Flight No. 1 80

TABLE VI Telemetry and Measurements of Bio-Flight No. 1 Measurements - Telemetered* Frequency Response Electrocardiogram 1-2 millivolts input 160 cps Heart sounds 600 cps Respiration thermistor 10/second Body temperature Thermistor (90° - 110°F) 10/second Ambient temperature thermistor (15o - 40°C) 10/second Ambient pressure (0-18 psi) 10/second *Also recorded on electromagnetic tape throughout re-entry. Measurements - Physical Method Heavy nuclei cosmic rays Nuclear track plates Atmospheric composition Chemical analysis of absorbents and air Measurements - Pre- and Post-Launch Physiological laboratory evaluation TABLE VII Payload Characteristics of Bio-Flight No. 1 Total Volume, Internal 670 Cu. In. Total weight 29. 5 lbs. Over-all dimensions, internal 9. 75x12. 5x6. 75 in. Volume of mobilebead 12.5 cu. in. Volume of baralyme 35 cu. in. Oxygen supply 22 cu. in. at 1800 psi Volume of ©2 (std. cond. ) 44. 15 liters Insulation . 5 in fiberglas - . 5 in rubber Heat source 8. 5 w Temperature control Thermostat 55° - 65°F. Circulation Convection Missile connection 1-13 pin elec. connector Time of installation 5 1/2 hrs. prior to takeoff Electric requirements 1. 5 watts Electrical source 28 v nose cone batteries 81

TABLE VIU Physiological Factors and Telemetry of Bio-Flight No. (FM/FM: Carrier Frequency 239. 5 me) Meas. No. Variable RDB RDB Sub-Carrier Channel Response Frequency 12A12 EKG B 5-14 330 cps 22 kc 12A13 Respiration B 5A-1 10 per sec. 30 kc 12A14 12A15 Body Temperature B 5A-2 10 per sec. 30 kc 12A16 Ambient Temperature B 5A-3 10 per sec. 30 kc 12A17 Ambient Pressure B 5A-4 10 per sec. 30 kc 12A18 Ambient Temperature A 5A-5 10 per sec. 30 kc 12A19 Ambient Pressure A 5A-6 10 per sec. 30 kc 12A20 Relative Humidity A 5A-7 10 per sec. 30 kc 12A21 Pulse Velocity A1 5-6 25 cps 1. 7 kc (12A23) 12A22 Percentage CC>2 5A-9 10 per sec. 30 kc 12A23 Pulse Velocity ^ 5-6 25 cps 1. 7 kc (12A21) 12A24 Respiration A 5A-13 10 per sec. 30 kc 12A25 Body Temperature A 5A-14 10 per sec. 30 kc 12A26 EKG A 5-13 220 cps 14. 5 kc 12A27 Behavioral Response A 5A-15 10 per sec. 30 kc 12A28 Reinforcement Shock A 5 A- 16 10 per sec. 30 kc 12A29 EMG1 A 5-17 790 cps 52. 5 kc 12A30 EMG2 A 5-16 600 cps 40 kc 12A31 Heart Sound A 5-18 1050 cps 70 kc 8. Cost of experiment depending on type of mission (non- interference with prime mission or biological experiment mission of flight), system adaptation and checkout, modification and inte- gration of payload into missile system, electronics, etc. May ap- proach price of missile if major alterations are necessary. 9. Number of recovery missions unknown. 10. Bibliography, including classified reports, is included in the general section at the end of this report. 82

1--- . f * Figure 57. Oscillographs of Bio- Flight No. 2 THOR and THOR-ABLE Combinations 1. During the first half of 1958, a series of ICBM re-entry nose cone tests was performed, using as the launching vehicle a two-stage missile consisting of a THOR IRBM and the Aerojet 1040 liquid-propellant rocket. This program, known as Project Able, was carried out by the Air Force and Space Technology Labora- tories, and included Project MIA: i. e. , Mouse-In-Able. It was followed by the Discoverer Project, a series of recoverable satel- lite shots, established on 28 February 1958 under the authority of ARPA. A staff of about 30 scientists and professionals is engaged in the preparations of experiments, the primary objectives of which include the exploration of extended periods of zero-G on animal behavior. The shots are prepared and launched at the Pacific Missile Range. As of November 1, 1960, 16 Discoverer shots have been made, several with animals. Two capsules have been recovered from 83

(US ABMA) Figure 58. Interior of Bio-Capsule for Squirrel Monkey (US ABMA) Figure 59. Electronic Package for Squirrel Monkey Capsule 84

(US ABMA) Figure 60. Bio.-Capsule for Rhesus Monkey Installed in JUPITER Nose Cone orbit, neither with animals. On August 21, 1960 the Soviets re- covered two dogs, 40 mice, and a variety of other organisms and tissues from a 24 hour orbit ("Sputnik 5"). On October 13, 1960, 3 black mice were recovered by the USAF from a 650 mile high, 5000 mile range flight in the RVX-2A 2800 pound reentry vehicle, boosted by an Atlas. 2. THOR is powered by a 150, 000 lb. thrust liquid pro- pellant engine (Rocketdyne) burning liquid oxygen and RP-1 high grade kerosene. In addition to the gimbaling action of the sustainer, thrust adjustment and roll control are obtained by two vernier engines located at the base of the missile on each side of the sus- tainer. The inertial guidance system is completely self-contained within the missile and cannot be deterred from a preset course. The nose cone is built by General Electric Company. They are protected by heat shields of ablation material, some of which have been recovered after re-entry shots over 1,500 miles at 10,000 miles per hour. THOR-ABLE combinations carried a bio-pack of about 20 pounds to a maximum altitude of 1,400 miles over a range of almost 5,500 miles. 3. Total flight duration of THOR about 15 minutes; of THOR- ABLE, about 45 minutes. Maximum acceleration liftoff about 12 g for THOR; about 5 g for second stage (ABLE). Period of virtual weightlessness in THOR about 500 seconds; in THOR-ABLE about 85

43 minutes. Peak decelerations at re-entry of THOR about 40 g; of THOR-ABLE about 60 g. 4. Physical conditions of specimen similar to those in the JUPITER IRBM, but less space in nose cone. Project MIA con- sisted of a mouse in a sealed capsule having a self-regenerating artificial atmosphere flown in a THOR-ABLE combination. A telemetered record of the heart beat was obtained intermittently throughout two of three ABLE flights (Figure 61). 1000 a oo 600 400 200 0 ZOO 400 600 100 1000 IZOO 1400 1600 ISOO ZOOO 22OO 2400 2600 2100 TIME FROM LIFT-OFF (SECONDS) -_.>HEART RATE '",•,ALTITUDE LASKA> HEART RATE L*SK*««ALTITUDE (After van der Wall and Young: ARS Journ.. Oct. 1959) Figure 61. Heart Rate and Altitude During Weightlessness Able Nos. 2 & 3 5. The animal was harnessed so that it could turn about a pivot point. The electrodes were implanted and attached beneath the skin. In designing the electronic system the mouse was con- sidered as an electrical generator having the following character- istics: open circuits volts, 1.5 mv peak-to-peak; impedance, 5,000 ohms; frequency range, 3 to 18 pulses per second. A pair of two-stage transistorized amplifiers was included in the unit and an attenuator was used to allow for variations in mouse output signal amplitude. Each amplifier circuit included about 10 db of negative feedback. This increased the input impedance to match the mouse impedance and stabilize the amplifier gain versus environmental changes. 6. Subsequent to amplification, silicon diodes rectified the heart beat pulses and an RC filter in the output circuit elongated them to increase the low-frequency component. It was necessary to avoid multi-vibrators because of their instability, although they would have increased the low frequency pulse rate component. Also, in order to prevent power supply decoupling networks, separate batteries were employed for the two amplifiers so that 86

only under very improbable circumstances could they oscillate. The millivolt-level mouse output pulses were transformed into 5-volt signals with a time-constant of about 50 milli-seconds, throughout the 4-day period of useful battery life. No evidence of adverse environmental effects on the electronics was detected in the records obtained from the flight. 7. Safety features and requirements were according to range safety regulations. Since the project was not classified as a hold item. Zero minus 16 hours was established as MIA-unit ready time. This preflight time, including possible holds during count- down, flight time, anticipated recovery period, and safety factor led to the design of a system which was to function properly for at least 72 hours. 8. Data on costs available at STL. 9. Experiments with the Mark I and Mark II biomedical re- covery capsule were scheduled through 1960-61. Experiments will include rhesus monkeys with life support system weighing up to 300 pounds. Recovery is to be accomplished through parachute and air snatch technique. Less than 5 G will be experienced by speci- men on air snatch. Surface vessels will back up air recovery operations. Six biomedical vans are available at Vandenberg AFB for ground support of this project. A biomedical experiment in Discoverer III, launched from Vandenburg Air Force Base on 4 June 1959 was only partly suc- cessful because of failure of the vehicle to gain sufficient velocity. However, part of the biomedical mission objectives were achieved. The life support system of the Mark I biomedical recovery capsule, developed by the AFBMD Directorate in coordination with the U.S. Air Force School of Aviation Medicine, functioned satisfactorily from liftoff to 790 seconds. The capsule atmosphere, continuously moved through a ducted gas control system, varied between 330 to 360 mm Hg oxygen partial pressure, maintained a low pCO2 and low relative humidity (below 60%) with a constant but relatively low temperature of 56°F. During flight, no leaks occurred in the cap- sule or oxygen system and the latter functioned normally to main- tain a satisfactory cell pressure between 6. 3 to 7. 0 psia from an oxygen cylinder maintaining approximately 1200 psig. Four C-57 black mice were specimens for the Discoverer III flight. An analysis of telemetry records revealed that all four animals lived through the burning stage and free-flight period until final loss of telemetry at 795 seconds. The animals were 87

active, particularly during the 500 seconds of weightlessness. Fre- quently, all four specimens showed movement simultaneously in response to vibration, tumbling, noise, and similar effects. Respi- ratory and EKG patterns could not be differentiated because of amplifier oscillation in all channels. However, on one viability signal, 29 seconds of EKG were reproduced from one specimen during the ballistic trajectory, and marked activity was recorded during weightlessness. The experience gained from this experi- ment seems to justify the enormous effort expanded on the Dis- coverer III launch and indicates the probability of success of sub- sequent bio-satellites of the Mark I and Mark II Discoverer series. 10. Discoverer III Biomedical Data Report. Directorate of Bioastronautics Projects, AFBMD, Hq, ARDC, WDZPB Report No. 2. ATLAS. AGENA and CENTAUR 1. ATLAS, AGENA and CENTAUR are some of the space vehicles to be employed for research purposes within the National Aeronautics and Space Administration program. Of these, only the ATLAS is available at this date. 2. The ATLAS (SM-65) ICBM having a range of about 5. 500 miles is produced by Convair Division of General Dynamics Corp- oration, San Diego, California, which builds the airframe and integrates the subsystems; General Electric Company, Philadelphia, Pennsylvania, and Syracuse, New York, manufacture the nose cone and the Radio Guidance System; Burroughs Company produces the computer; American Machine and Foundry furnished the power supply accessories; and North American Aviation, Rocketdyne Division, manufactures the propulsion system. The ATLAS, which has been undergoing flight tests since June, 1957, at the Air Force Missile Test Center, Atlantic Missile Range, Cape Canaveral, Florida, begins its flight with all three engines in operation developing about 400, 000 pounds of thrust. Several seconds after liftoff, the booster engines drop away and the missile continues with the lower thrust sustainer, reaching ap- proximately 15,000 mph maximum. Two small vernier rockets located on the lower portion of the 10-foot body supply roll control and thrust adjustment throughout the powered flight. The ATLAS will be used as booster in the second phase of the Discoverer and of the Mercury project. 88

It is difficult to make an exact estimate of the payload weights which will be carried by any of these vehicles on the various missions. By and large, it can be assumed that these larger systems can put in earth-near orbits payloads which amount to about 1% of their liftoff weight. This ratio becomes more favor- able with the bigger boosters. In all cases, the advanced space program is primarily aimed at the exploration of the moon, the planets, and interplanetary space, and is not intended to directly contribute to the Man-in-Space and other bioastronautical projects. However, test flights will most probably precede the actual as- signed missions of the vehicles; and the Saturn will be employed for a moon-miss development test. There exists a certain possibility that biological payloads could be carried on some of the earlier sub-orbital and orbital flights, and that some of them could be re- covered after ejection from the upper stages. 3. Guidance and attitude-control systems will be employed to control the trajectories of these vehicles. During the launch phase, injection guidance will be used, but radio and inertial guidance seems to be more desirable for the upper stages. The development of these systems will rely upon developments which have already been undertaken in guidance systems for military rockets. The first vehicles will be equipped with devices which are the same as, or similar to, those already available. As the projects accelerate, lighter weight and more advanced guidance systems designed specifically for the space vehicle application will be employed. 4. Not available at this time. 5. Not available at this time. 6. The magnetic tape recorder will be the primary data- storage tool in the next years. Flaws in telemetry, particularly due to wow and flutter and to signal drop out, will have to be overcome in order to improve this type of frequency-modulated telemetry. The maximum frequency which can be recorded today is in the order of 100 kc with a tape speed of 60 in/sec. By uti- lizing new techniques of recording and amplification, an upper limit of 500 kc at 60 in/sec, or 250 kc at 30 in/sec tape speed, can be obtained. Future developments in the area of field data presentation will be primarily directed toward the improvement of analog data plotting on a direct-writing oscillograph, and toward the addition of the digital data to the analog record in order to provide one record containing good-accuracy analog with high- accuracy digital data. 89

7. According to range safety and vehicle requirements. 8. Not available. 9. Not available for sub-orbital and satellite flights. 10. "Exploration of the Moon, the Planets, and Interplanetary Space", (Edit, by Albert R. Hibbs). Report No. 30-1. Jet Pro- pulsion Laboratory, California Institute of Technology, Pasadena, California, April 30, 1959 (Unclassified). LITTLE JOE Vehicle 1. The LITTLE JOE rocket is used as a research rocket by the National Aeronautics and Space Administration. Its vehicle airframes were designed and constructed according to NASA speci- fications by the Missile Division of North American Aviation, Inc. Motors were built and supplied by the Thiokol Chemical Corpora- tion. The propulsion system consists of a cluster of eight solid propellant rocket engines which can be made up of either four Pollux or Castor main rocket motors and four Recruit auxiliary rocket motors. The four Recruit rocket engines and two of the main rocket engines are fired simultaneously at launch. The re- maining rockets are fired together at some pre-determined time after launch. Two biomedical experiments have been done so far with LITTLE JOE rockets at Wallops Island, Virginia. 2. A sketch of the LITTLE JOE vehicle is shown in Figure 62. This vehicle was built specifically to obtain flight performance data on Project MERCURY research and development type cap- sules. It is conceivable that a variety of payloads could be de- signed and flown on the LITTLE JOE booster system. Photographs of the LITTLE JOE booster and capsule payloads before and after mating are shown in Figures 63 and 64. 3. Motion, acceleration, and velocity characteristics are classified and can be obtained from the NASA Space Task Group, Langley Air Force Base, Virginia. By and large, periods of weightlessness of several minutes can be obtained, depending upon payload weight and flight trajectory. 4. The gross weight of the capsule payload shown in Figures 62, 63, and 64 was about 3,000 pounds. The interior capsule pressure was 1 atmosphere at 100% oxygen at the start of the 90

Figure 62. Schematic Model of LITTLE JOE Vehicle experiment and reduced to 350 mm Hg through breathing during flight. The capsule temperature was kept between 10° and 20°C in both flights. The NASA capsule can be position stabilized through the use of reaction control jets which operate on compressed gas. 5. The LITTLE JOE booster is a fin-stabilized vehicle and employs no guidance and control system. A method of compen- sating for winds which affect the flight path has been developed, utilizing adjustment of launch azimuth and elevation, and has been used successfully. 6. Typically instrumented NASA capsule used on LITTLE JOE boosters provide from 50 to 100 channels of information depending on the nature of the test. Some measurements are recorded on- board the capsule, others pass through a telemeter link. The fol- lowing list of instrumentation is a sample list for one of the NASA LITTLE JOE capsules. The "T only" notation indicates measure- ments that were telemetered only. The items with just a "T" are recorded onboard and telemetered. All other items were recorded onboard. The number of instruments is indicated in parentheses. 91

Accelerometers Rate gyros Internal pressures Capsule air temperature Voltages Frame counters GSAP Bio-pack channels a and 0 vanes Manifold pressure monitor External pressures Total pressure Up camera Side camera Down camera Onboard timer Noise pickups Impact accelerometers Drogue chute tensiometer Capsule Marman band release (3) (3) (2) (1) (2) (3) (11) (2) (1) (9) (1) (1) (1) (2) (1) (2) (3) (1) (1) "T only" "T only" (NASA) Figure 63. LITTLE JOE Booster and Capsule Before Mating 92

(NASA) Figure 64. LITTLE JOE Research Vehicle Capsule separation (1) Escape tower Marman band release (1) Drogue cover separation (1) Main chute deploy (1) Camera in bio-pack (1) "T" llrr-lll It rpl I llrptl The data from LITTLE JOE firings are reduced by the NASA and suppliers of special test packages. 7. All circuitry on the LITTLE JOE vehicle is carefully checked before and during countdown. Nearly constant monitoring of important circuits or instrumentation can be provided during the countdown. In case of an emergency the booster is equipped with a command destruct system. The destruct system is actually a sys- tem for terminating the thrust of the main rocket motors. 93

8. The approximate cost of a LITTLE JOE booster is $500,000. Payload costs would, of course, be dependent upon the size and complexity of the payload. 9. The LITTLE JOE program is basically a program to test the escape capabilities of a MERCURY capsule at the maximum aerodynamic loading conditions that might occur during an escape off a MERCURY-ATLAS. Four LITTLE JOE boosters with their capsule payloads have been launched as of this date. All four flights have been termed successful with regard to booster per- formance. The two biomedical experiments, using rhesus monkeys, done by scientists from the School of Aerospace Medicine, Brooks Air Force Base, Texas, involved physiological instrumentation, psy- chological performance, and environmental measurements. The flight profile of the rockets included maximum accelerations of about 10 - 12 g, and weightlessness period of about 3 minutes at plus or minus . 02 G. The peak altitude obtained in the 3rd flight was about 300,000 feet. Measurements taken concerned the EKG, respiration, eye movements, and bar pressing behavior, but only partial results were obtained due to malfunction of equipment. The attempt to obtain inflight pictures during the first experiment only partially successful because of malfunction of the camera equipment. Moreover, O2 tension, total pressure, capsule temp- erature, and relative humidity were recorded. The second animal experiment (January 1960) was done in conjunction with an engi- neering flight test, reaching a maximum altitude of about six miles. 10. NASA Space Task Group: Information Concerning LITTLE JOE (Confidential report specially prepared for use in "Zero-G Devices and Weightlessness Simulators"). Further information can be obtained from the Director of the Space Task Group, NASA, Langley Field, Virginia. SATURN 1. The SATURN development program calls for initial ve- hicle of two or three stages, depending on specific mission. It will be about 200 feet high in the three stage configuration. NASA assumed technical direction over SATURN under a NASA-DOD agreement of 18 November 1959. The SATURN booster project was initiated in August, 1958. The first booster flight with dummy upper stages is scheduled for the 1961-1962 time frame. 94

The objective of Project SATURN is to provide by the 1963- 1964 time period an efficient and reliable system for lifting multi- ton loads into high orbit around the earth and into deep space. It will be for general use in the heavy space vehicle field. First stage of the SATURN vehicle will be a super-booster with 1,500,000 pounds of thrust, plus upper stages. The super- booster is under development at Marshall Space Flight Center. It can put a payload of 15 tons into a low satellite orbit, about 5 tons around the moon, and about one ton softly on the surface of the moon. Two men can make a lunar circumnavigation with sub- sequent return to the earth. If a refueling maneuver is carried out in a low satellite orbit, a SATURN vehicle will be capable of landing two men on the moon, andof bringing them back home a few days later. In short, SATURN will be the first American- made vehicle that opens the lunar space to man. Pictures of the SATURN system, lunar landing vehicles, and a proposed orbital station to be developed as part of this system are shown in Figures 65 and 65 a-d. Figure 65. SATURN Booster During Static Firing 2. The SATURN booster uses a grouping of engines built from existing JUPITER and THOR components. The cluster con- cept was chosen because of the proven reliability of the existing engine. The H-1 engine is considerably more compact than com- panion JUPITER and THOR engines in the same 150, 000-pound thrust range, resulting in a major move for simplification and high reliability. The first of these engines was delivered to 95

A >\ (NASA Marshall Space Flight Center) Figure 65a. Orbital Launched Lunar Return Vehicle AOMC on May 1, 1959. The booster unit is being designed, fabri- cated, and tested at the Missile Agency. Rocketdyne Division of North American Aviation is supplying the basic H-1 engines. The individual H-1 engines used in the propulsion system develop thrust from the propulsive gases created by the com- bustion of liquid oxygen and RP-1, a hydro-carbon fuel similar to that used in turbo-jet engines. Combustion within the thrust chamber occurs at temperatures above 5000 degrees Fahrenheit. The walls of the thrust chambers are formed of thin, lightweight metal through which fuel is forced at high speed. The fuel flow prevents the metal from melting under the extremely high tempera- tures encountered. Lightweight turbopumps, proven in repeated flights of JUPITER and other ballistic missiles, are utilized in the H-l engine to draw the propellants from the storage tanks and force them at high flow rate into the combustion chamber. 96

The booster is powered by a cluster of eight Rocketdyne H-1 engines using liquid oxygen, and RP-1 propellants contained in 9 tanks. Total thrust of the clustered engines will be equivalent to 30,000,000 horsepower. The first stage cluster of H-1 engines will be about 80 feet tall, 22 feet in diameter, and will produce about 1,500,000 pounds of thrust at liftoff. The vehicle will use an all-inertial guidance system. This system will automatically compensate for deviations resulting from loss of thrust should one of the first stage booster engines fail to function. The first stage cluster was static tested on a special 177- foot tower at Marshall Space Flight Center in the spring of 1960. Flight tests will be made from the Atlantic Missile Range at Cape Canaveral, Florida, where a 245-foot self-propelled service stand is being built to handle the huge rocket. MANNED ORBITAL TRANSPORT (NASA Marshall Space Flight Center) Figure 65b. Manned Orbital Transport 97

(NASA Marshall Space Flight Center) Figure 65c. Lunar Landing Vehicle Directional stability of the entire missile will be maintained by swiveling the four outer engines of the cluster. These engines will be mounted on gimbals--a part of their assembly—and moved by struts actuated by the guidance equipment of the space vehicle. 3. The SATURN program in general offers the following capabilities in the area of weightlessness and space missions: a. The earliest and most economical possibility for surpassing all presently known payload capa- bilities for orbital and trans-orbital research flights. b. Earliest capability for large orbital space stations and bioastronautical laboratories. c. Earliest capability for manned lunar circumnaviga- tion and return. 98

d. Earliest capability for manned lunar landing and return. e. Earliest available booster for research on large winged and manned space vehicles. The acceleration and deceleration stresses expected to occur with the SATURN system vehicles are within the range of human tolerance. However, we are dealing with relatively low accelera- tions over a considerable period of time, or higher accelerations (decelerations) of relatively short durations. In either case, it will be necessary to furnish adequate protection against the physio- logical effects, and particularly to provide for an adequate safety factor in case of emergency. Whatever the size of the capsule or cabin may be, we assume that the passenger will be protected either by seats or couches similar to those used in the MERCURY capsule, or by dynamic positioning during peak acceleration. Moreover, high G suits and similar devices may be used during take-off, orbital maneuver as- sociated with high accelerations, and decelerations during return and landing. 1 i 1,1 i / \ LUNAR LANDING (NASA Marshall Space Flight Center) Figure 65d. Lunar Earth Return Vehicle 99

It seems almost unnecessary to say that the state of weight- lessness will prevail during all periods of non-accelerated or free flights. 4. The theoretical basis and equations representing the phy- sical conditions for the subject have been given before. A repre- sentation of the various trajectories and orbits associated with booster capabilities and payload penalties cannot be given in this context, nor is it possible to cite the physical parameters and equations pertaining to orbital and trans-orbital flights. 5. Control system variables concern payload attitude con- trolling and control loop in which the subject is the human link. The control of the attitude of the payload may be required for: a. Physiological reasons (tolerance of the passenger to angular acceleration associated with spinning and tumbling). b. Position and orientation maneuvering (guidance observa- tion). c. Radio communication and navigation. d. Scientific measurements, including photography, aiming of solar cells at the sun, taking fixes, measurements, etc. The function of an attitude control system is to maintain prescribed relationships between reference directions in the vehicle and an external frame of reference. The attitude sensing device may include radiation detectors, such as photo cells, and inertial elements, such as gyroscopes. This latter type is limited by its inability to maintain a reference direction over a long period of time. On the other hand, the radiation detector can maintain a reference direction over an indefinite period of time. Further- more, accurate determination of direction is possible with optical devices, particularly if the vehicle is manned. A distinction must be drawn between attitude control systems intended only to maintain the angular attitude of the vehicle, and systems which are capable of varying and controlling the attitude in accordance with commands. Systems utilizing spin stabilization or differential gravity effects belong to the first class. Command 100

signals of a more complex nature will be used for most space missions, at least during mid-course corrections and terminal guidance. The accuracy of these systems may be greatly in- creased if human navigational aid is supplementary. The role of the human passenger in the SATURN space vehicle system must be considered from two aspects. The first concerns his proficiency as part of the control link. Careful consideration must be given to his control and performance parameters so that he is protected against the possible hazards of space and his pro- ficiency is maintained. For this, he must be furnished adequate and reliable information under the zero-G condition. Preparatory investigation of zero-G human engineering by means of weightless flights in rocket vehicles described before will furnish preliminary information. Secondly, the effects of weightlessness on the human subject must be determined, which are needed for the improvement of control link design adequate for zero-G conditions. Since large payloads can be put in orbit, all sorts of experiments on zero-G effects can be carried out. This includes the testing of the space cabin, the engineered environment, and the adequacy or practica- bility of space cockpit designs. The immediate goal is the develop- ment and testing of an advanced, integrated, and reliable cabin and life support system which proves to be adequate under weightless conditions of actual orbital flight. 6. The collection and processing of biomedical information in a SATURN type space vehicle will not pose any serious problems. The use of magnetic recorders for this purpose has been dealt with in the description of the ATLAS, AGENA and CENTAUR class. In addition, a world-wide system of ground stations will be available for bio-telemetry. The amount and type of data which are pre- sented in the individual field stations will depend greatly upon the requirements for real-time data analysis. Generally speaking, any type of presentation needed in the final data reduction can be made. Improvements in existing communication facilities will make it possible to receive tracking signals and scientific information from a distance of about two million miles. The areas in which improvements will be made to obtain greater communication ranges concern energy sources, miniaturization, antennas, and amplifier systems. The weight required to transmit vehicle and scientific information is proportional to the expression 2 / KTMLR2 ' * * dt (57) EDAtAr 101

where: K = Boltzmann's constant T = effective temperature of the receiver detector M = sampling rate of the experiment L = effective signal-to-noise ratio desired from the receiver R = distance of information transmission E = effective efficiency of transmission system D = storage capacity of batteries A^A = effective areas of the transmitting and receiving antennas, respectively, and the integral is taken over the time of transmission. The product KTMLR2 is directly proportional to the power out- put of the system. Each field station available for the recording of scientific data will contain adequate display such as a direct writing os- cillograph and digital print-out systems. With good design and proper calibrations, data from these records are read with greater accuracy than the oscillograph analog method provides. The ac- curacy of the digital data is limited by the data frequency, counter- sampling, frequency of the subcarrier channel, and the accuracy limit of the counter. However, the present systems used for the mass accumula- tion of scientific information are inadequate for long-range measure- ments. At present, electrical signals representing physiological or physical quantities are registered continuously in real-time at the output of these systems. Each signal source requires an individual channel, monopolizing transmission, and final display facilities regardless of signal activity or information content. The equipment bandwidth and the weight and power penalties are fixed in accordance with the maximum potential demand by signals in all channels ir- respective of the normal or average signal value. With the signi- ficant signals averaging less than 2% of the total, the over-all efficiency of the system is notoriously low. The same flood of signals imposes serious problems at the output end because of 102

equipment limitations on the one hand, and because of the restricted capacity of the human operators on the other hand. Thus, the read-out becomes a real problem when days, weeks, and continuous operations are involved. In order to remedy this situation, McClennan (WADC) has proposed a "viability monitor" which would discriminate between the flood of insignificant data and the significant bits of informa- tion. The technique suggested assumes that every biomedical measurement can be defined in such a way as to allow an electronic decision whether a bit of information is significant or otherwise. Such a device would include a normalizer, a gating unit, and a decision unit. In this circuit, the standardized voltage for gating is normally supplied from the decision unit. The decision to pass signals is made by amplitude coincidence of the signal voltage with a fixed reference voltage. In order not to lose significant infor- mation, a preselector may be applied to "alert" the decision unit after long periods of insignificant flow of information. By employing modern automatic data reduction methods, the telemetry measurements can be rapidly and efficiently reduced to a presentable form. To accomplish this, a data handling net is necessary in conjunction with a world-wide tracking system. Since many field stations located around the globe will receive signals only part of the time, a central data processing center must be established. After the data are reduced at the acquisition site, they will be transmitted to the central headquarters. A global radio communication system will be available for this purpose. In summary, it is believed that signals selection, storage, and burst transmittal (the latter by command or automatically) will be applicable for biotelemetry in SATURN space vehicles. On the other hand, normal data collection and evaluation within the vehicle will also be feasible because of the relatively high pay- load allowances for orbital flights. 7. The SATURN engines are profiting by the ten-year development period on the REDSTONE, JUPITER, THOR, and ATLAS engines. Continual analysis of the static test firing re- ports on these engines have permitted design of the H-1 on the basis of elimination of trouble spots, and improvements of other components. Each of the eight H-1 engines comprising the SATURN cluster operates independently with its own gas generator and turbo- pump. The propellant tankage supplying liquid oxygen and RP-1 fuel is a common source for all engines. The start sequence is the simplest of any yet provided in the family of liquid propellant 103

engines. Safety devices have been built into the engine--an added factor in manned flight. One of the safety factors is the assurance that all eight engines of the cluster must be functioning normally before the missile can be released. Hypergolic ignition (self- igniting) will further increase the reliability of the H-1. The potential use of the SATURN vehicle for manned space flight does not necessarily increase the reliability of the system. However, there is a strong belief that reliability of the planned missions can be increased substantially through the participation of a human crew. Another effort must be made to obtain a manned rating for the SATURN through test and experience, even though the basic design features and requirements for the safety of human occupants have been incorporated early in the design of SATURN vehicles. 8. The SATURN system promises to become the most eco- nomical means for space research and transportation in the fore- seeable future. The cost for orbital flight will be reduced to ap- proximately $500 per pound or less in the SATURN program. More- over, booster recovery seems to offer a potential saving. A special system employing parachutes and retro-rockets is being developed by the Cook Electric Company, Morton Grove, Illinois, to enable recovery of the first stage booster for post-flight inspection and possible further use. 9. NASA's space science efforts in terms of its present program concern various scientific programs within the areas of planetary space sciences, astronomy, and astrophysics. A life science program was developed in 1960, which includes the use of SATURN vehicles after they are available. In general, weightlessness research in the SATURN type vehicle will continue where the limitations of the research aircraft made further investigations impossible. The following tentative areas appear to be worth-while for further investigation: a. Basic biological study. Research concerns the develop- ment of life under zero-G conditions, and the processes of growth, maturation, fertilization, reproduction, aging, and decay of living matter. Specimens for this type of research are single cells and cell systems, tissue, plants, and animals of a relatively low phylogenetic standing. Related subjects are the adjustment of lower organisms to the zero-G condition, geotropism, and the effect of other environmental factors (temperature, pressure, C2, CC>2, radiation, nutrition, etc.) upon the weightless organism. 104

b. Studies on the consequences of relativity theory and re- lated subjects on the organism. These investigations concern possible mass changes due to time dilation, including alteration of growth rates and aging. Although it appears doubtful that such effects may be detected with this type of research vehicle, pre- liminary or feasibility studies of these problems may be possible. Moreover, the effect of differential gravitational forces during orbital and trans-orbital flights, and possible differences between the zero-G and the true agravic state upon organic matter and the human organism, will be investigated. c. Human tolerances and adaptation to long exposure to weightlessness and readjustment to the 1 g condition after prolonged flights. These studies include (1) clinical investigation of certain critical physiological functions (e.g., cardiodynamics, hemo- dynamics, respiration, digestion, hormone production, etc. ) under conditions of extended weightlessness, (2) disturbances of sensory perception and neuromusclar condition during prolonged states of weightlessness, (3) the possible occurrence of muscular atrophy due to inactivity of the skeletal muscles, (4) perception, orienta- tion, and performance characteristics, and (5) the possible phy- siological and psychological stress of extended periods of weight- lessness, the effect of drugs and medication upon tolerance, and preventative measures and exercise. This also includes preventa- tive means for readjustment of the astronauts to the normal gravi- tational condition. d. Human engineering parameters in extended periods of weightlessness. The main subject of research in this area must be to develop devices, instruments, means, and other hardware which will help the astronaut to live, work, and relax in the weight- less state. This mainly concerns the manufacturing of seats, beds, tools, and equipment adequate for the life in a weightless world. Since our instruments and devices were developed on the principles of the normal force of gravity, every tool and device which depends upon the attraction of the earth or is constructed according to the open loop principle must be redesigned for the mechanics of weightlessness. It is important that only the mass of a body be used in all calculations, for mass--and not weight—is constant under these conditions. Moreover, the application of accelera- tions to produce an artificial gravity and its effect upon man will be studied. e. Selection of space crews. Flights in the SATURN vehicle will be used to test human tolerances to weightlessness and to select astronauts and space crews. 105

f. Training of space crews. It is believed that the SATURN vehicle can be used for the training of personnel with regard to adjustment to and performance in the weightless condition. g. Research and psychological consequences of prolonged states of weightlessness. There is a certain possibility that the physical conditions which personnel of an artificial satellite or in a space ship will have to endure may have detrimental effects upon the emotional condition. Since many of the habit patterns, which were unconsciously established over the years, will suddenly be brought to an awareness, mental effort will have to be employed to do what formerly was automatically performed. It is in these conditions that man must work, rest, and sleep. The possibility of traumatic experiences and the prevention of emotional dis- turbances will be subjects of investigation. h. Studies of the effects of subgravity. Lunar landings and stay on the surface of the moon will permit the investigation of the effects of subgravity (about one-sixth of the terrestrial gravity) on plants, animals, and man. 10. References: a. ABMA Report No. DSP-TM-13-59, (U) SATURN System Study II, CONFIDENTIAL report, dated 13 November 1959. b. ABMA Report No. DSP-TM-1-59, (U) SATURN System Study, SECRET report, by H. H. Koelle, F. L. Williams, and W. G. Huber, dated 13 March 1959. c. AOMC SECRET report, (U) Supplement No. 2 to Development and Funding Plan for ARPA Order 14-59, dated 24 July 1959. d. ABMA Report No. DIR-TR, 2-59, (U) Semi-Annual Technical Summary Report on ARPA Orders 14-59 and 47-59, SECRET report, dated 13 July 1959. e. (U) Final Report - Economical Boost Systems Study, MD 59-147, CONFIDENTIAL report, Missile Division, North American Aviation, Inc. , Downey, California, dated 2 October 1959. f. (U) Missile Road Transport Survey, M-M-M1-59-53, Unclassified report, by L. M. St. Ours, Martin Company, Denver, Colorado, dated September 1959. g. JPL Report No. 30-1, Exploration of the Moon, the Planets, and Interplanetary Space, Unclassified report, dated 30 April 1959. 1Q6

H. WEIGHTLESSNESS SIMULATORS In the search for methods of studying the effects of weight- lessness on the ground, immersion in water has attracted con- siderable interest. According to the Archimedian Principle, a body which is partly or wholly immersed in a fluid loses its weight by an amount equal to that of the displaced medium: W = V U-«j' ) g, (58) where V is the volume and S and S represent the densities of the body and the surrounding fluid, respectively. The apparent weight Wa of a body, being equal to the sum of the forces of gravity and inertia, can be expressed as: wa =Fg +Fi =- Fs- (59) where Fg, Fj, and Fs represent the forces of gravity, inertia, and support, respectively. It is evident from (58) and (59) that a body can be made weightless not only by counterbalancing Fg through FI (free-fall), but also by placing the body in a fluid of the same density (i. e. , by making 6 = S' ). Theoretically, the density of the body can be larger, equal to, or smaller than that of the fluid. Accordingly, the weight of the immersed body can be positive, zero, or negative, respectively. The body, being a three-dimensional and extended object, is at least partly compressed by the hydrostatic pressure of the surrounding fluid. If the body is submerged at greater depth, this pressure is considerable, no matter whether the weight of the displaced fluid is smaller or larger than that of the body (note that the numerical value of this weight differs from the apparent weight described above, although the concepts, logically and operationally, are identical). Within the body, gravitation produces a hydrostatic effect, too; i. e. , the pressure which a certain column of fluid (such as blood) exerts on the walls of the vessel. This pressure is no longer exerted if the body is immersed in water, because it is counter- acted by the hydrostatic pressure of the fluid surrounding the body. Since no particular surface area of the body has to carry weight and the exteroceptors of the skin are neutralized because of the pressure equilibrium inside and outside of the body, the latter 107